Abstract:
A turbine vane usable in a turbine engine and having at least one cooling system. The cooling system may include at least one convergent flow channel for receiving air from a shroud assembly. The cooling system may also include impingement channels in a leading edge cavity for impinging a cooling fluid against an inner surface of a leading edge of the turbine vane. The cooling system may also include a serpentine cooling path for removing heat from aft sections of the turbine vane proximate to the trailing edge of the turbine vane. The cooling system may also include a divergent leading edge cavity. Exterior film cooling is not needed to safely operate a turbine vane according to this invention.

Description:
FIELD OF THE INVENTION  
       [0001]     This invention is directed generally to turbine vanes, and more particularly to hollow turbine vanes having cooling channels for passing cooling fluids, such as air, to cool the vanes and supply cooling fluids to the manifold of a turbine assembly.  
       BACKGROUND  
       [0002]     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 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 prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.  
         [0003]     Typically, turbine vanes are formed from an elongated portion forming a vane having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to a manifold. The vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. The 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 aspects of the turbine vane at a relatively uniform temperature. At least some of the air passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the vane. A substantially portion of the air is passed into a manifold to which the vane is movable coupled. The air supplied to the manifold may be used, among other uses, to cool turbine blade assemblies coupled to the manifold. While advances have been made in the cooling systems in turbine vanes, a need still exists for a turbine vane having increased cooling efficiency for dissipating heat and passing a sufficient amount of cooling air through the vane and into the manifold.  
       SUMMARY OF THE INVENTION  
       [0004]     This invention relates to a turbine vane having a cooling system including a convergent flow channel for receiving cooling fluids from a shroud assembly and passing a portion of the cooling fluids to one or more impingement channels in a leading edge cooling cavity and allowing the remainder of the cooling fluids to pass through a serpentine cooling path before being exhausted through exhaust orifices in the trailing edge of the turbine vane. The cooling system has the capacity to sufficiently cool the turbine vane without requiring external film cooling orifices.  
         [0005]     The turbine vane may be formed from a generally elongated hollow airfoil having a leading edge, a trailing edge, a pressure side, a suction side, a first end adapted to be coupled to a shroud assembly, and a second end opposite the first end adapted to be coupled to a manifold assembly. The convergent flow channel may include an inlet generally at the first end of the airfoil and may extend toward the second end of the airfoil. The convergent flow channel may have a first cross-sectional area proximate to the first end of the airfoil that is larger than a second cross-sectional area of the convergent flow channel closer to the second end of the airfoil than the first cross-sectional area. This configuration of the convergent flow channel enables the cooling system to regulate flow of cooling fluids into the manifold assembly and to prevent overheating of the trailing edge of the vane.  
         [0006]     The turbine vane may also include a plurality of impingement channels extending from the convergent flow channel toward the leading edge and terminating in a leading edge cooling cavity aft of an inner surface of the leading edge in a leading edge cooling cavity. The impingement channels may vary in length such that a first channel located closest to the first end of the airfoil may be shorter than a second impingement channel closest to the second end of the airfoil. In at least one embodiment, each impingement channel may terminate at a substantially similar distance from the inner surface of the leading edge to maintain high impingement jet velocity and high impingement cooling effectiveness. This configuration is achieved by increasing the length of each impingement channel moving from the first end of the airfoil to the second end of the airfoil. The cross-sectional area of each impingement channel may be substantially equal or may vary. Likewise, the distance between each impingement channel may be substantially equal or may vary as well.  
         [0007]     In at least one embodiment, one or more of the plurality of impingement channels may be positioned within the leading edge cooling cavity using one or more pin fins. The pins fins may extend from an inner surface of the suction side of the vane and attach to an impingement channel or may extend from the inner surface of the pressure side of the vane and attach to the impingement channel, or both. In at least one embodiment, each of the impingement channels is held in position using pin fins. The pin fins increase the surface area available for convection, thereby increasing the cooling capacity of the cooling system.  
         [0008]     In at least one embodiment, the convergent flow path forms a portion of a serpentine cooling path in an aft portion of the turbine vane. The serpentine cooling path may have numerous passes, which in at least one embodiment may number three passes. The serpentine cooling path may be in communication with one or more exhaust orifices in the trailing edge of the turbine vane for exhausting cooling fluids from the vane.  
         [0009]     In operation, a cooling fluid enters the cooling system from a shroud assembly through one or more inlets in the first end of the turbine vane. The cooling fluid enters the convergent flow channel and, a substantial portion of the cooling fluid is then bled off of the convergent flow channel through the impingement channels. The cooling fluid flows through the impingement channels and impinges against the inner surface of the leading edge. The cooling fluid then flows through the leading edge cooling cavity and is exhausted to the manifold assembly. The cooling fluids remaining in the convergent flow channel is passed through a serpentine cooling path and exhausted through one or more exhaust orifices in the trailing edge of the blade.  
         [0010]     An advantage of this invention is that the cooling system is capable of removing sufficient heat without necessitating external film cooling.  
         [0011]     Another advantage of this invention is that the leading edge cooling cavity may be configured as a divergent cooling cavity, which minimizes cross flow of the cooling fluids passing through impingement channels proximate to the first end of the airfoil.  
         [0012]     Yet another advantage of this invention is that the pin fins increase the cooling capacity of the cooling system.  
         [0013]     These and other embodiments are described in more detail below. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0014]     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.  
         [0015]      FIG. 1  is a perspective view of a turbine vane having features according to the instant invention.  
         [0016]      FIG. 2  is cross-sectional view of the turbine vane shown in  FIG. 1  taken along line  2 - 2 .  
         [0017]      FIG. 3  is a cross-sectional view of the turbine vane shown in  FIGS. 1 and 2  taken along line  3 - 3  in  FIG. 2 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0018]     As shown in  FIGS. 1-3 , this invention is directed to a turbine vane  10  having a cooling system  12  in inner aspects of the turbine vane  10  for use in turbine engines. The cooling system  12  is configured such that adequate cooling occurs internally without using external film cooling from cooling fluids supplied through orifices in the housing forming the vane  10 . In particular, the cooling system  12  includes at least one convergent flow channel  14  for receiving a cooling fluid from a shroud assembly  16 , and may include one or more impingement channels  18  proximate to a leading edge  20  for directing cooling fluids to contact an inner surface  22  of the leading edge  20 . In at least one embodiment, the convergent flow channel  14  may be a serpentine cooling path  24 , which directs a cooling fluid through one or more exhaust orifices  26  in a trailing edge  28  of the turbine vane  10 .  
         [0019]     As shown in  FIG. 1 , the turbine vane  10  may be formed from a generally elongated airfoil  30  having an outer surface  32  adapted for use in an axial flow turbine engine. Outer surface  32  may be formed from a housing  34  having a generally concave shaped portion forming pressure side  36  and may have a generally convex shaped portion forming suction side  38 . The turbine vane  10  may also include a first end  40  adapted to be coupled to the shroud assembly  16  and a second end  42  adapted to be coupled to a manifold assembly  44 .  
         [0020]     As shown in  FIG. 2 , the convergent flow channel  14  may have a first cross-sectional area  46  proximate to the first end  40  of the airfoil  30  that is larger than a second cross-sectional area  48  closer to the second end  42  of the airfoil  30  than the first cross-sectional area  46 . In at least one embodiment, the convergent flow channel  14  may extend from the first end  40  of the airfoil  30  to a second end  42  of the airfoil  22 . In other embodiments, the convergent flow channel  14  may not extend the entire length between the first and second ends  40 ,  42 . In at least one embodiment, the convergent flow channel  14  may be a first inflow section  52  of the serpentine cooling path  24 . The serpentine cooling path  24  may also include a first outflow section  54  and a second inflow section  56  forming a three-pass serpentine cooling path. The serpentine cooling path  24  is not limited to a three-pass system, but may have additional or fewer flow paths. Exhaust orifices  26  may be positioned in the trailing edge  28  and provide a pathway for cooling fluids to be exhausted from the second inflow section  56 . In at least one embodiment, the serpentine cooling path  24  may include trip strips  64  for mixing cooling fluids as the cooling fluids flow through the serpentine cooling path  24  and for increasing the amount of heat removed from the turbine vane  10 .  
         [0021]     The convergent flow channel  14  may be formed from at least one rib  50  positioned between the leading edge  20  and the convergent flow channel  14 . The rib  50  may be positioned in a generally nonparallel position relative to the leading edge  20 , which forms a divergent leading edge cooling cavity  68 . The divergent leading edge cavity  68  receives cooling fluids from the impingement channels  18 . The divergent leading edge cooling cavity  68  minimizes the cross flow effect of cooling fluids flowing parallel to the inner surface  22  of the leading edge  20  and thereby, maximizes heat transfer at the inner surface  22 . The rib  50  may include one or more orifices  51  to which the impingement channels  18  may be coupled. In at least one embodiment, as shown in  FIG. 2 , the rib  50  may include a plurality of orifices  51  to which impingement channels  18  may be coupled. One or more impingement channels  18  may extend from the rib  50  to towards an inner surface  22  of the leading edge  20 . In at least one embodiment, the impingement channels  18  may terminate in the divergent leading edge cooling cavity  68  aft of the inner surface  22  of the leading edge  20 . Each impingement channel  18  may terminate at a substantially equal distance from the inner surface  22  of the leading edge  20 , which allows cooling fluids flowing through the impingement channels  18  to maintain a high impingement jet velocity and impingement cooling effectiveness. The impingement channels  18  may have substantially equal cross-sectional areas or may have cross-sectional areas having difference sizes. The impingement channels  18  may be spaced apart at substantially similar distances or at equal distances.  
         [0022]     In at least one embodiment, as shown in  FIG. 2 , the turbine vane  10  may include a plurality of impingement channels  18  extending between the rib  50  and the leading edge  20  and positioned from the first end  40  of the airfoil  30  to the second end  42  of the airfoil  30 . The impingement channels  18  regulate the flow of cooling fluids through the turbine vane  10  and prevent overflow of cooling fluids to the manifold assembly  44 . By preventing overflow to the manifold assembly  44 , the possibility of overheating portions of the housing  34  proximate to the trailing edge  28  is reduced. The impingement channel  18  positioned at the first end  40  may have the shortest length of the impingement channels  18  positioned between the first and second ends  40 ,  42 . The impingement channels  18  may increase in length proceeding from the first end  40  to the second end  42 . In other words, each impingement channel  18  may be longer than the impingement channel  18  immediately adjacent to the channel  18  and closer to the first end  40  of the airfoil  30 . The impingement channels  18  may be positioned at a substantially equal distance from each other or may be positioned a varying distances from each other.  
         [0023]     In at least one embodiment, the impingement channels  18  may be held in position between an inner surface  58  of the suction side  38  and an inner surface  60  of the pressure side  36  using one or more pin fins  62 . One or more of the impingement channels  18  may be supported by a pin fin  62  positioned between an inner surface  60  of the pressure side  36  and the impingement channel  18 , or positioned between an inner surface  58  of the suction side  38  and the impingement channel  18 , or both. The pin fins  62  increase the surface area of the housing  34  and thereby increase the amount of convection surfaces.  
         [0024]     In operation, a cooling fluid enters the cooling system  12  through an inlet  66  in the convergent flow channel  14 . The inlet  66  may be sized and configured to regulate the flow of cooling fluids into the convergent flow channel  14 . The cooling fluids are bled into the impingement channels  18  from the convergent flow channel  14 . The cooling fluids flow through the impingement channels  18  and are exhausted into the leading edge cool cavity  68 . The cooling fluids impinge against the inner surface  22  of the leading edge  20 . The cooling fluids then flow through the leading edge cooling cavity  68  to the manifold assembly  44 . In at least one embodiment including a divergent leading edge cooling cavity  68 , the negative effects of cooling fluid cross flow is reduced to the point of being almost negligible because the cavity  68  increases in cross-sectional area as additional cooling fluid is emitted from each impingement channel  18 , moving from the first end  40  to the second end  42  of the airfoil  30 . Thus, cross-flow velocity is maintained at a substantially steady rate. Cooling fluids not flowing into the impingement channels  18  continue to flow through the serpentine cooling path  24  and are exhausted through the exhaust orifices  26 . The amount of cooling fluids flowing through the turbine vane  10  and into the manifold assembly  44  is controlled by the number and cross-sectional areas of the impingement channels  18 .  
         [0025]     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.