Abstract:
The present application provides a method of installing an impingement cooling assembly in an inner platform of an airfoil of a turbine nozzle. The method may include the steps of positioning an insert within a cavity of the airfoil, positioning a core exit cover about an opening of the cavity, positioning an impingement plenum within a platform cavity, inserting an unfixed spoolie through an assembly port of the impingement plenum and into an airflow cavity of the insert, and closing the assembly port.

Description:
TECHNICAL FIELD 
       [0001]    The present application and the resultant patent relate generally to gas turbine engines and more particularly relate to methods for assembling cooling components in an inner platform of a cantilevered turbine nozzle and the like with reduced leakage. 
       BACKGROUND OF THE INVENTION 
       [0002]    Impingement cooling systems have been used with turbine machinery to cool various types of components such as casings, buckets, nozzles, and the like. Impingement cooling systems cool the components via the airflow so as to maintain adequate clearances between the components and to promote adequate component lifetime. One issue with some types of known impingement cooling systems, however, is that they tend to require complicated casting and/or structural welding. Such structures may not be durable or may be expensive to produce and repair. Moreover, the components required for impingement cooling should be tolerant of manufacturing variations and tolerant of thermal differentials between, for example, the nozzle vanes, the shrouds, the sheet metal, the plumbing hardware, and other components. These tolerance requirements may result in significant gaps between the components so as to cause undesirable leakage between pressure cavities. 
         [0003]    There is thus a desire for tightly packaged cooling components for use with turbine nozzles and methods of assembling the same. Preferably the cooling components may allow the nozzle to adequately face high gas path temperatures while meeting lifetime and maintenance requirements as well as being reasonable in cost. Moreover, assembly of these components may be simplified and reduce any gaps therebetween that may lead to leakages. 
       SUMMARY OF THE INVENTION 
       [0004]    The present application and the resultant patent provide a method of installing an impingement cooling assembly in an inner platform of an airfoil of a turbine nozzle. The method may include the steps of positioning an insert within a cavity of the airfoil, positioning a core exit cover about an opening of the cavity, positioning an impingement plenum within a platform cavity, inserting an unfixed spoolie through an assembly port of the impingement plenum and into an airflow cavity of the insert, and closing the assembly port. 
         [0005]    The present application and the resultant patent further provide an impingement cooling assembly for use in an inner platform of a turbine nozzle. The impingement cooling assembly may include an impingement insert positioned about an airfoil cavity of the nozzle, an impingement plenum with an assembly port positioned about the inner platform and the impingement insert, and a spoolie extending from the impingement plenum about the assembly port and into the airfoil cavity of the nozzle. 
         [0006]    These and other features and improvements of the present application and the resultant patent 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 
         [0007]      FIG. 1  is a schematic diagram of a gas turbine engine showing a compressor, a combustor, and a turbine. 
           [0008]      FIG. 2  is a partial side view of a nozzle vane with an impingement cooling assembly therein. 
           [0009]      FIG. 3  is an exploded view of a nozzle vane with an impingement cooling assembly as may be described herein. 
           [0010]      FIG. 4  is a partial section view of the nozzle vane with the impingement cooling assembly of  FIG. 3 . 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    Referring now to the drawings, in which like numerals refer to like elements throughout the several views,  FIG. 1  shows a schematic view of gas turbine engine  10  as may be used herein. The gas turbine engine  10  may include a compressor  15 . The compressor  15  compresses an incoming flow of air  20 . The compressor  15  delivers the compressed flow of air  20  to a combustor  25 . The combustor  25  mixes the compressed flow of air  20  with a pressurized flow of fuel  30  and ignites the mixture to create a flow of combustion gases  35 . Although only a single combustor  25  is shown, the gas turbine engine  10  may include any number of combustors  25 . The flow of combustion gases  35  is in turn delivered to a turbine  40 . The flow of combustion gases  35  drives the turbine  40  so as to produce mechanical work. The mechanical work produced in the turbine  40  drives the compressor  15  via a shaft  45  and an external load  50  such as an electrical generator and the like. 
         [0012]    The gas turbine engine  10  may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine  10  may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine  10  may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together. 
         [0013]      FIG. 2  is an example of a nozzle  55  that may be used with the turbine  40  described above. Generally described, the nozzle  55  may include a nozzle vane  60  that extends between an inner platform  65  and an outer platform  70 . A number of the nozzles  55  may be combined into a circumferential array to form a stage with a number of rotor blades (not shown). 
         [0014]    The nozzle  55  also may include an impingement cooling assembly  85  with an impingement plenum  90 . The impingement plenum  90  may have a number of impingement apertures  95  formed therein. The impingement plenum  90  may be in communication with the flow of air  20  from the compressor  15  or another source via a spoolie or other type of cooling conduit. The flow of air  20  may extend through the nozzle vane  60 , into the impingement cooling assembly  85 , and out via the impingement apertures  95  so as to impingement cool a portion of the nozzle  55  or elsewhere. Other components and other configurations may be used herein. 
         [0015]      FIG. 3  and  FIG. 4  show portions of an example of a nozzle  100  as may be described herein. In this example, a multivaned segment  110  is shown with a first vane  120  and a second vane  130 . Any number of vanes and any number of segments may be used herein. The vanes  120 ,  130  may extend from an inner platform  140 . The inner platform  140  may a platform cavity  160 . Each of the vanes  120 ,  130  may include an airflow cavity  170  therein. The airflow cavity  170  may be in communication with the platform cavity  160  so as to provide the flow of air  20  from the compressor  15  or elsewhere for impingement cooling. Other components and other configurations may be used herein. 
         [0016]    The nozzle  100  also may include an impingement cooling assembly  180  therein. The impingement cooling assembly  180  may include an impingement plenum  190 . The impingement plenum  190  may include one or more spoolies or other types of cooling conduits in communication with the flow of air  20  from the airflow cavities  170 . The spoolies or conduits may include both coolant passages and housings designed to minimize gaps with interfacing components. In this configuration, a first spoolie  200  and a second spoolie  210  are shown. Any number of spoolies may be used. In this configuration, the first spoolie  200  may be positioned in a first housing  300  and the second spoolie  210  may be positioned in a second housing  310 . The nozzle  100  may also include a number of airfoil sheet metal inserts. In this configuration, a first insert  230  may be contained within the first vane  120  and a second insert  250  may be contained within the second vane  130 . A core exit cover may be affixed to the exit of each vane cavity. In the current configuration, a first core exit cover  220  may be affixed to an opening  225  of the first vane  120  and a second core exit cover  240  may be affixed to an opening  245  of the second vane  130 . The impingement plenum  190  also may include the assembly port  260 , an assembly port cover  270 , and a retention plate  280 . The current example shows a single assembly port and assembly port cover but multiples may be used of each. The impingement plenum  190  and the components thereof may have any size or shape. Other components and other configurations may be used herein. 
         [0017]    In order to assemble the impingement cooling assembly  180 , the airfoil inserts  230 ,  250  may be positioned within the airfoil cavities  170 . The core exit covers  220 ,  240  may be welded or otherwise affixed into place. The impingement plenum  190  may be fabricated with the first spoolie  200  welded or otherwise affixed into place. The impingement plenum  190  may be positioned within the platform cavity  160  such that the first spoolie  200  engages the first airfoil insert  230 . The second spoolie  210  may be positioned within the assembly port  260  and into engagement with the second airfoil insert  250 . The assembly port  260  may be sized to accommodate the spoolies passing therethrough with sufficient provision for alignment of the spoolie with the airfoil insert to minimize the hydraulic gaps between the components. The second spoolie  210  may be welded or otherwise affixed to the impingement plenum  190 . The assembly port cover  270  then may be welded or otherwise affixed into place about the assembly port  260 . Additional cover plates also may be used. Multiple assembly ports may be used with all of the spoolies being positioned into engagement with airfoil inserts through the assembly ports prior to being affixed to the impingement plenum  190 . 
         [0018]    The retention plate  280  then may be slid into place circumferentially. The retention plate  280  may take the form of a seal carrier  290  and the like. The retention plate  280  may be held in place via a retention pin or other types of mechanical engagement. Other components, such as seals or gaskets, also may be used herein. Other configurations may be used herein. The order of the installation and assembly steps herein may vary. The impingement cooling assembly  180  thus is assembled from the inner diameter outward. 
         [0019]    The impingement cooling assembly  180 , and the methods described herein, thus may minimize hydraulic gaps between cavities of differing pressures. Specifically, the methods may minimize cross-cavity leakage while remaining tolerant of manufacturing variations. The impingement cooling assembly  180  may be mechanically retained without complex welding or castings. Lower leakage thus equates to higher overall performance and efficiency. 
         [0020]    It should be apparent that the foregoing relates only to certain embodiments of the present application and the resultant patent. 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.