Abstract:
A cooling system for a turbine airfoil of a turbine engine having multiple segmented ribs aligned together spanwise within a trailing edge cooling channel. The segmented ribs may be positioned proximate to a trailing edge of the turbine airfoil to facilitate increased heat removal with less cooling fluid flow, thereby resulting in increased cooling system efficiency, and to increase the structural integrity of the trailing edge of the airfoil. The segmented ribs may include crossover orifices that provide structural integrity to ceramic cores used during manufacturing to prevent cracking and other damage.

Description:
FIELD OF THE INVENTION 
     This invention is directed generally to turbine airfoils, and more particularly to cooling systems in hollow turbine airfoils. 
     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 having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade 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. 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. Often, conventional turbine blades develop hot spots in the trailing edge of the blade. While the trailing edge of the turbine blade is not exposed to as harsh of conditions as a leading edge of the blade, the trailing edge requires cooling nonetheless. Thus, a need exists for a cooling system capable of providing sufficient cooling to composite airfoils while also providing sufficient structural support to the airfoil as well. 
     SUMMARY OF THE INVENTION 
     This invention relates to a turbine airfoil cooling system including a trailing edge cooling channel with at least one segmented rib having a plurality of impingement orifices. The segmented rib increases the efficiency of the cooling system in the airfoil and increases the strength of the airfoil in the trailing edge region. The trailing edge cooling channel may be configured such that during manufacturing of the channel, the likelihood of damage to a ceramic core used to create the internal cooling channels is reduced. The trailing edge cooling channel may be configured such that a ceramic core used to produce the airfoil has greater structural strength, thereby reducing the risk of cracking and other damage to the ceramic core during formation of the airfoil. 
     The turbine airfoil may be formed from a generally elongated airfoil having a leading edge, a trailing edge, a tip section at a first end, a root coupled to the airfoil at an end generally opposite the first end for supporting the airfoil and for coupling the airfoil to a disc, and at least one cavity forming a cooling system in the airfoil. The turbine airfoil may include at least one trailing edge cooling channel extending from the root to the tip section of the elongated airfoil. The trailing edge cooling channel may include at least one first segmented spanwise rib positioned in the at least one trailing edge cooling channel, that extends generally from the root to the tip section of the elongated airfoil. The first segmented spanwise rib may include a plurality of impingement orifices. 
     The trailing edge cooling channel may also include at least one second segmented spanwise rib positioned in the at least one trailing edge cooling channel that extends generally from the root to the tip section of the elongated airfoil. The trailing edge cooling channel may be positioned between the first segmented spanwise rib and the trailing edge of the generally elongated airfoil and include a plurality of impingement orifices. In another embodiment, the trailing edge cooling channel may include at least one third segmented spanwise rib extending generally from the root to the tip section of the elongated airfoil and positioned in the at least one trailing edge cooling channel between the second segmented spanwise rib and the trailing edge of the generally elongated airfoil. The third segmented spanwise rib may also include a plurality of impingement orifices. 
     The plurality of impingement orifices may increase turbulence in the trailing edge cooling channel, thereby increasing the effectiveness of the cooling channel by increasing the convection rate in the channel. In at least one embodiment, the impingement orifices in the segmented ribs may be offset spanwise relative to the impingement orifices in upstream segmented ribs. 
     In another embodiment, the segmented cooling channels may include crossover orifices that provide a cooling fluid pathway through the segmented cooling channels and structural integrity to a ceramic core used to produce the cooling channel. In at least one embodiment, crossover orifices may be positioned between ends of the segmented cooling channels and the tip section and between an opposite end of the segmented cooling channels and the root. Such a configuration enables a rectangular support structure to be formed within a ceramic core used to create the airfoil with an internal cooling channel. The rectangular support structure greatly enhances the structural integrity of the ceramic core in the trailing edge region, thereby reducing the likelihood of damage to the ceramic core during the manufacturing process. 
     The crossover orifices in the adjacent segmented ribs may be aligned spanwise. Alternatively, the crossover orifices may be offset spanwise in the adjacent segmented ribs. In yet another embodiment, the segmented ribs may not include cross-over orifices. The crossover orifices may be distinguishable from the impingement orifices in that the crossover orifices may have a cross-sectional area that is greater than a cross-sectional area for the impingement orifices. In at least one embodiment, the crossover orifices may have a cross-sectional diameter generally equal to a distance between inner surfaces of the suction and pressure sides. 
     During use, cooling fluids, which may be, but are not limited to, air, flow into the cooling system from the root of the airfoil. At least a portion of the cooling fluids flow into the trailing edge cooling channel. The cooling fluids flow spanwise through the impingement orifices in the segmented ribs. In embodiments in which the impingement orifices and the crossover orifices are offset, cooling fluids pass through a rib and impinge on a downstream rib. The cooling fluids increase in temperature, thereby reducing the temperature of the airfoil. The cooling fluids are discharged through either orifices or through trailing edge orifices. 
     An advantage of this invention is that the segmented ribs form a rectangular grid structure that increase the ceramic core stiffness, thereby minimizing the likelihood of ceramic core breakage during manufacturing and improving the manufacture cast yields. 
     Another advantage of this invention is that the segmented ribs increase the cross-sectional area of the ceramic core at the ribs, which reduces the risk of core breakage due to shear forces developed from differential shrink rates of the ceramic core, external shell and molten metal. 
     Yet another advantage of this invention is that the increased cross-sectional area of the core of the airfoil increases the moment of inertia, which in turn improves the resistance to local edge bending at the trailing edge and total bending at the trailing edge. 
     Another advantage of this invention is that the invention improves ceramic core breakage modes, such as shear, local edge bending, and overall bending at the trailing edge, thereby creating a stiffer trailing edge for a ceramic core during manufacturing with reduced risk of breakage due to overall trailing edge bending and improved manufacturability. 
     These and other embodiments are described in more detail below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are incorporated in and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention. 
         FIG. 1  is a perspective view of a turbine airfoil having features according to the instant invention. 
         FIG. 2  is cross-sectional view, referred to as a filleted view, of the turbine airfoil shown in  FIG. 1  taken along line  2 - 2 . 
         FIG. 3  is cross-sectional view of the turbine airfoil shown in  FIG. 2  taken from the line  3 - 3  in  FIG. 2 . 
         FIG. 4  is a cross-sectional view of an alternative embodiment of the turbine airfoil shown in  FIG. 1  taken from the same perspective as line  2 - 2 . 
         FIG. 5  is cross-sectional view of the turbine airfoil shown in  FIG. 4  taken from the line  5 - 5 . 
         FIG. 6  is a cross-sectional view of an alternative embodiment of the turbine airfoil shown in  FIG. 1  taken from the same perspective as line  2 - 2 . 
         FIG. 7  is cross-sectional view of the turbine airfoil shown in  FIG. 6  taken from the line  7 - 7 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     As shown in  FIGS. 1-7 , this invention is directed to a turbine airfoil cooling system  10  for turbine airfoil  12  used in turbine engines. In particular, the turbine airfoil cooling system  10  is directed to a cooling system  10  located in a cavity  14 , as shown in  FIGS. 2-7 , positioned between two or more walls  18  forming a housing  16  of the turbine airfoil  12 . The cooling system  10  may include a trailing edge cooling channel  20  adapted to receive cooling fluids to reduce the temperature of the turbine airfoil  12  thereby reducing the required cooling fluid flow to achieve adequate cooling and increasing the effectiveness of the cooling system  10 . The trailing edge cooling channel  20  may be configured such that during manufacturing of the channel  20 , the likelihood of damage to a ceramic core is reduced. The trailing edge cooling channel  20  may be configured such that a ceramic core  68 , which forms the cavities  14 , used to produce the airfoil  12  has greater structural strength, thereby reducing the risk of cracking and other damage during formation of the airfoil  12 . 
     The turbine airfoil  12  may be formed from a generally elongated airfoil  22  coupled to a root  24  at a platform  26 . The turbine airfoil  12  may be formed from conventional metals or other acceptable materials. The generally elongated airfoil  22  may extend from the root  24  to a tip section  36  and include a leading edge  34  and trailing edge  38 . Airfoil  22  may have an outer wall  18  adapted for use, for example, in a first stage of an axial flow turbine engine. Outer wall  18  may form a generally concave shaped portion forming pressure side  28  and may form a generally convex shaped portion forming suction side  30 . The cavity  14 , as shown in  FIGS. 2-7 , may be positioned in inner aspects of the airfoil  22  for directing one or more gases, which may include air received from a compressor (not shown), through the airfoil  22  and out one or more orifices  32  in the airfoil  22  to reduce the temperature of the airfoil  22  and provide film cooling to the outer wall  18 . The cavity  14  may include trip strips  70 , as shown in  FIGS. 2 ,  4 , and  6 . The trip strips  70  may be positioned nonparallel to the direction of flow of the cooling fluids through the cavity  14 . As shown in  FIG. 1 , the orifices  32  may be positioned in a leading edge  34 , a tip section  36 , or outer wall  18 , or any combination thereof, and have various configurations. The cavity  14  may be arranged in various configurations and is not limited to a particular flow path. 
     The cooling system  10 , as shown in  FIGS. 2-7 , may include a trailing edge cooling channel  20  for removing heat from the airfoil  22  proximate to the trailing edge  38 . The trailing edge cooling channel  20  may include one or more segmented ribs  40  extending generally spanwise within the cooling channel  20 . In at least one embodiment, the segmented ribs  40  extend generally from the root  24  to the tip section  36 . However, in an alternative embodiment, the segmented ribs  40  may be formed in other lengths. As shown in  FIGS. 2 ,  4 , and  6 , the trailing edge cooling channel may be formed from a first segmented rib  42 , a second segmented rib  44 , and a third segmented rib  46 . The segmented ribs  42 ,  44 ,  46  may extend generally spanwise and parallel to each other. The third segmented rib  46  may be positioned between the second segmented rib  44  and the trailing edge  38  of the airfoil  22 , and the second segmented rib  44  may be positioned between the first segmented rib  42  and the trailing edge  38 . The segmented ribs  42 ,  44 ,  46  may extend from the pressure side  28  to the suction side  30 . In alternative embodiments, the trailing edge cooling channel may include greater than or fewer than three segmented ribs. 
     The segmented ribs  42 ,  44 ,  46  may include one or more impingement orifices  48 . The impingement orifices  48  may be sized, such as those shown in  FIGS. 3 and 7 , to have a diameter that is smaller than a distance between an inner surface  50  of the pressure side  28  and an inner surface  52  of the suction side  30 . The impingement orifices  48  may also have a substantially hourglass cross-sectional shape in which an inlet  54  tapers to a smaller diameter center region  58 . Similarly, an outlet  56  may taper to the center region  58  as well. Alternatively, the impingement orifices  48  may have other appropriate sizes. The impingement orifices  48  may be offset relative to each other. For instance, as shown in  FIG. 2 ,  4 ,  6 , the impingement orifices  48  in the second segmented rib  44  may be offset relative to the impingement orifices  48  in the first segmented rib  42 . Similarly, the impingement orifices  48  in the second segmented rib  44  may be offset relative to the impingement orifices  48  in the third segmented rib  42 . Offsetting the impingement orifices  48  creates convection rate increasing turbulence in the trailing edge cooling channel  20  by causing cooling fluids to impinge on downstream segmented ribs  40 . 
     The segmented ribs  42 ,  44 ,  46  may include one or more crossover orifices  60  that break the ribs  42 ,  44 ,  46  into a plurality of parallel, aligned segments  62 . The crossover orifices  60  provide structural integrity to a ceramic core  68  used to manufacture the airfoil  12 . The crossover orifices  60  may be larger in cross-sectional area than the impingement orifices  48 . In at least one embodiment, as shown in the embodiments in  FIGS. 3 ,  5 , the crossover orifices  60  may extend from the inner surface  50  on the pressure side  28  to the inner surface  52  on the suction side  30 . The crossover orifices  60  may have other sizes as well. 
     The segmented ribs  42 ,  44 ,  46  may include one or more crossover orifices  60  along their lengths. In at least one embodiment, the crossover orifices  60  may be positioned between the ribs  42 ,  44 ,  46  and the tip section  36  and between the ribs  42 ,  44 ,  46  and the root  24 . Such a configuration forms a generally rectangular support structure in a ceramic core  68  used to form the trailing edge cooling channel  20 . The rectangle extends along the trailing edge  38  of the airfoil  12 , along the tip section  36  and the root  24 , and the portion of the ceramic core  68  used to form the cavity  14  proximate to the first segmented rib  42 . The rectangular support structure greatly improves the reliability of the ceramic core  68  while reducing the risk of cracking and damage to the ceramic core  68  before the ceramic core  68  is removed later in the manufacturing process through conventional leaching processes. 
     The crossover orifices  60  may be aligned spanwise, as shown in  FIG. 4 . Alternatively, the crossover orifices  60  may be offset from each other. For example, as shown in  FIG. 2 , the crossover orifices  60  in the third segmented rib  46  may be offset spanwise from the crossover orifices  60  in the second segmented rib  44 . Similarly, the crossover orifices  60  in the second segmented rib  44  may be offset spanwise from the crossover orifices  60  in the first segmented rib  42 . In yet another embodiment, the segmented ribs  42 ,  44 , and  46  may not include any crossover orifices  60 , but include only impingement orifices  48 . 
     The trailing edge cooling channel  20  may also include a plurality of support ribs  66  positioned in close proximity to the trailing edge  38 , as shown in  FIGS. 2 ,  4 , and  6 . The support ribs  66  may have any configuration appropriate for increasing the strength of the airfoil  22  to reduce local trailing edge bending and overall trailing edge bending. In the embodiments shown in  FIGS. 2 ,  4 , and  6 , the support ribs  66  may have a generally rounded upstream corner and conclude at the trailing edge  38 . 
     During operation, cooling fluids, which may be, but are not limited to, air, flow into the cooling system  10  from the root  24 . At least a portion of the cooling fluids flow into the cavity  14  and into the trailing edge cooling channel  20 . The cooling fluids flow spanwise through the impingement orifices  48  in the segmented ribs  42 ,  44 ,  46 . In embodiments in which the impingement orifices  48  and the crossover orifices  60  are offset, cooling fluids pass through a rib  42 ,  44 ,  46  and impinge on a downstream rib  44 ,  46 . The cooling fluids increase in temperature, thereby reducing the temperature of the airfoil  22 . The cooling fluids are discharged through either orifices  32  or through trailing edge orifices  64 . 
     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.