Abstract:
The present application thus provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a seal positioned adjacent to the turbine component. The seal may include a number of apertures so as to allow the airflow to pass therethrough.

Description:
TECHNICAL FIELD 
       [0001]    The present application relates generally to gas turbine engines and more particularly relates through cooling hot engine components via airflow through seal components. 
       BACKGROUND OF THE INVENTION 
       [0002]    A portion of the airflow through a turbine engine may be diverted and used for cooling purposes. The overall efficiency of the turbine engine, however, is decreased by the amount of air that is diverted for cooling purposes as opposed to being used for combustion. The less airflow diverted for cooling purposes or other types of parasitic airflows, the better the efficiency and operation of the gas turbine engine as a whole. 
         [0003]    By way of example, a retainer ring positioned about a stage one nozzle outer base and a stage one shroud may be cooled by a compressor discharge airflow. The retainer ring may include a number of circumferential grooves and radial slots therein. The retainer ring may be cooled with a cooling flow from the core airflow. Circumferential cooling of the retaining ring, however, may cause a circumferential temperature gradient therein. Moreover, machining the grooves in the retaining ring requires machining time and labor costs. 
         [0004]    There is thus a desire for improved systems and methods for component cooling that involves less airflow while increasing overall system efficiency. The systems and methods preferably also allow the use of less complicated and costly components 
       SUMMARY OF THE INVENTION 
       [0005]    The present application thus provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a seal positioned adjacent to the turbine component. The seal may include a number of apertures so as to allow the airflow to pass therethrough. 
         [0006]    The present application further provides for a turbine component cooling system. The component cooling system may include a turbine component, an airflow passing adjacent to the turbine component, and a W-seal positioned adjacent to the turbine component. The W-seal may include a number of apertures so as to allow the airflow to pass therethrough. 
         [0007]    The present application further provides for a turbine component cooling system. The component cooling system may include a stage one nozzle outer base, a retaining ring positioned adjacent to the a stage one nozzle outer base, a stage one shroud positioned adjacent to the retaining ring, an airflow passing adjacent to the retaining ring along an air sealing plate, and a seal positioned between the air sealing plate and the stage one shroud. The seal may include a number of apertures so as to allow the airflow to pass therethrough. 
         [0008]    These and other features of the present application will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  is a side cross-sectional view of a gas turbine engine. 
           [0010]      FIG. 2  is a side cross-sectional view of the inner section of a stage one nozzle outer base and a stage one shroud showing a retaining ring. 
           [0011]      FIG. 3  is a side cross-sectional view of the inner section of a stage one nozzle outer base and a stage one shroud showing a retaining ring and a W-seal as is described herein. 
           [0012]      FIG. 4  is a side plan view of the W-seal as may be used herein. 
           [0013]      FIG. 5  is a cross-sectional view of the W-seal. 
       
    
    
     DETAILED DESCRIPTION 
       [0014]    Referring now to the drawings, in which like numbers refer to like elements throughout the several views,  FIG. 1  shows a cross-sectional view of a gas turbine engine  100 . As is known, the gas turbine engine  100  may include a compressor  110  to compress an incoming flow of air. The compressor  110  delivers the compressed flow of air to a combustor  120 . The combustor  120  mixes the compressed flow of air with a compressed flow of fuel and ignites the mixture. The hot combustion gases are in turn delivered to a turbine  130 . The hot combustion gases drive a number of turbine buckets  140  so as to produce mechanical work. The mechanical work produced in the turbine  130  drives the compressor  110  and also an external load such as an electrical generator and the like. The gas turbine engine  100  may use natural gas, various types of syngas, and other types of fuels. Other types of gas turbine engines also may be used herein. The gas turbine engine  10  may have other configurations and may use other types of components. Multiple gas turbine engines  100 , other types of turbines, and other types of power generation equipment may be used herein together. 
         [0015]      FIG. 2  is an expanded view of a portion of the turbine  130 . Specifically, the intersection of a retaining ring  145  as positioned about a stage one nozzle outer base  150  and a stage one shroud  160  as positioned within a turbine shell  170 . An air sealing plate  180  may be positioned between the retaining ring  145  and the stage one shroud  160 . The air sealing plate  180  faces a number of circumferential grooves  190  positioned within the retaining ring  145  on one side and one or more W-seals  200  on the other side. The retaining ring  145  with the air sealing plate  180  may be cooled by an airflow  195 , in this case a compressor discharge airflow  195  passing through the circumferential groves  190 . Such circumferential air cooling, however, may cause a temperature gradient along the retaining ring  145  of up to about  20  degrees Fahrenheit (about 6.7 degrees Celsius) or so. 
         [0016]      FIG. 3  shows a component cooling system  210  as is described herein. The component cooling system  210  may include a turbine component  220 . The turbine component  220  may be a retaining ring  225 . In this example, the retaining ring  225  includes an air sealing plate  230  but not the circumferential grooves  190 . The component cooling system  210  also may include one or more W-seals  240  positioned against the air sealing plate  210 . As is shown in  FIGS. 4 and 5 , the W-seal  240  may include a number of apertures  250  therein so as to permit a cooling airflow  255  therethrough. The apertures  250  may be about 0.07 inches (about 1.78 millimeters) in diameter and about 300 in number. The size and number of the apertures  250  may be varied with the desired cooling flow rate. 
         [0017]    The W-seal  240  may be made out of Inconel 718 or similar types of materials. (Inconel 718 is a nickel chromium alloy made precipitation hardenable by additions of aluminum and titanium and having creep rupture strength at high temperatures to about 1290 degrees Fahrenheit (about 700 degrees Celsius)). Inconel is a trademark of Huntington Alloys Corporation of Huntington, W.V. Other types or combinations of materials may be used herein. The Inconel material may have a thickness of about 0.01 inches (about 0.254 millimeters). 
         [0018]    The airflow  255  through the W-seal  240  may be about 0.215% W25. The use of the W-seals  240  may provide more uniform cooling of the retaining ring  225 . Material and labor costs in producing circumferential grooves  190  also are eliminated. Likewise, eliminating the circumferential grooves  190  makes the retaining ring  225  structurally stronger. The W-seal  240  thus both meters the cooling airflow  255  therethrough while acting as the seal between the retaining ring  145  and the stage one shroud  160 . The W-seals  240  described herein further may be used anywhere a cooling flow may be required. Other types of seals and other types of turbine components  220  may be used with the component cooling system  210  described herein. 
         [0019]    It should be apparent that the foregoing relates only to certain embodiments of the present application and that numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the invention as defined by the following claims and the equivalents thereof.