Abstract:
A method for repairing a turbine nozzle segment includes providing a turbine nozzle segment having an enclosure with an integral discourager, removing the enclosure and/or the discourager from the turbine nozzle segment and attaching a new enclosure ad/or discourager to the turbine nozzle segment.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The exemplary embodiments relate generally to gas turbine engine components and more particularly to turbine nozzle segments having improved cooling. 
         [0002]    Gas turbine engines typically include a compressor, a combustor, and at least one turbine. The compressor may compress air, which may be mixed with fuel and channeled to the combustor. The mixture may then be ignited for generating hot combustion gases, and the combustion gases may be channeled to the turbine. The turbine may extract energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. 
         [0003]    The turbine may include a stator assembly and a rotor assembly. The stator assembly may include a stationary nozzle assembly having a plurality of circumferentially spaced apart airfoils extending radially between inner and outer bands, which define a flow path for channeling combustion gases therethrough. Typically the airfoils and bands are formed into a plurality of segments, which may include one or two spaced apart airfoils radially extending between an inner and an outer band. The segments are joined together to form the nozzle assembly. 
         [0004]    The rotor assembly may be downstream of the stator assembly and may include a plurality of blades extending radially outward from a disk. Each rotor blade may include an airfoil, which may extend between a platform and a tip. Each rotor blade may also include a root that may extend below the platform and be received in a corresponding slot in the disk. Alternatively, the disk may be a blisk or bladed disk, which may alleviate the need for a root and the airfoil may extend directly from the disk. The rotor assembly may be bounded radially at the tip by a stationary annular shroud. The shrouds and platforms (or disk, in the case of a blisk) define a flow path for channeling the combustion gases therethrough. 
         [0005]    As gas temperatures rise due to the demand for increased performance, components may not be able to withstand the increased temperatures. Higher gas temperatures lead to higher metal temperatures, which is a primary contributor to distress. Bands are susceptible to this distress, in particular, on the aft side of the bands. Distress may cause cracking or holes to form within these areas, leading to decreased performance and higher repair costs. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    In one exemplary embodiment, a method for repairing a turbine nozzle segment includes providing a turbine nozzle segment having an enclosure with an integral discourager, removing the discourager from the turbine nozzle segment and attaching a new discourager to the turbine nozzle segment. 
         [0007]    In another exemplary embodiment, a method for repairing a turbine nozzle segment includes the steps of providing a turbine nozzle segment having an enclosure with an integral discourager, removing the enclosure and discourager from the turbine nozzle segment and attaching a new enclosure and discourager to the turbine nozzle segment. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0008]      FIG. 1  is a cross-sectional view of an exemplary gas turbine engine. 
           [0009]      FIG. 2  is a cross sectional view of an exemplary embodiment of a turbine nozzle assembly. 
           [0010]      FIG. 3  is close-up cross-sectional view of an exemplary embodiment of a band of a turbine nozzle assembly. 
           [0011]      FIG. 4  is a close-up plan view of the flow path side of an exemplary embodiment of a band of a turbine nozzle assembly. 
           [0012]      FIG. 5  is a prospective view of an exemplary embodiment of a turbine nozzle segment. 
           [0013]      FIG. 6  is another prospective view of an exemplary embodiment of a turbine nozzle segment shown with the enclosure removed. 
           [0014]      FIG. 7  is another prospective view of an exemplary embodiment of a turbine nozzle segment shown with the enclosure. 
           [0015]      FIG. 8  is a flow chart showing one exemplary embodiment of a repair method. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0016]      FIG. 1  illustrates a cross-sectional schematic view of an exemplary gas turbine engine  100 . The gas turbine engine  100  may include a low-pressure compressor  102 , a high-pressure compressor  104 , a combustor  106 , a high-pressure turbine  108 , and a low-pressure turbine  110 . The low-pressure compressor may be coupled to the low-pressure turbine through a shaft  112 . The high-pressure compressor  104  may be coupled to the high-pressure turbine  108  through a shaft  114 . In operation, air flows through the low-pressure compressor  102  and high-pressure compressor  104 . The highly compressed air is delivered to the combustor  106 , where it is mixed with a fuel and ignited to generate combustion gases. The combustion gases are channeled from the combustor  106  to drive the turbines  108  and  110 . The turbine  110  drives the low-pressure compressor  102  by way of shaft  112 . The turbine  108  drives the high-pressure compressor  104  by way of shaft  114 . 
         [0017]    As shown in  FIGS. 2-7 , the high-pressure turbine  108  may include a turbine nozzle assembly  116 . The turbine nozzle assembly  116  may be downstream of the combustor  106  or a row of turbine blades. The turbine nozzle assembly  116  includes an annular array of turbine nozzle segments  118 . A plurality of arcuate turbine nozzle segments  118  may be joined together to form the annular turbine nozzle assembly  116 . The turbine nozzle segments  118  may have an inner band  120  and an outer band  122 , which radially bound the flow of combustion gases through the turbine nozzle assembly  116 . The inner band  120  may have a flowpath side  124  and a non-flowpath side  126  and the outer band  122  may have a flowpath side  128  and a non-flowpath side  130 . One or more flanges  132  may extend from the non-flowpath sides  128  and  130  of the inner band  120  and outer band  122 . For example, as shown in  FIGS. 2 and 3 , flange  134  extends radially from said the inner band  120  and may be used to attach the turbine nozzle assembly  116  to other components of the gas turbine engine  100 . 
         [0018]    One or more airfoils  136  extend radially between the inner band  120  and outer band  122  for directing the flow of combustion gases through the turbine nozzle assembly  116 . The airfoils  136  have a leading edge  138  on the forward side of the turbine nozzle segment  118  and a trailing edge  140  on the aft side of the turbine nozzle segment  118 . The airfoils  136  may be formed of solid or hollow construction. Hollow airfoils may include one or more internal cooling passages for cooling the airfoil and providing film cooling to the airfoil surfaces. Other hollow airfoils may include one or more cavities for receiving a cooling insert. The cooling insert may have a plurality of cooling holes for impinging on the interior surface of the hollow airfoil before exiting as film cooling through holes in the airfoil. Any configuration of airfoil known in the art may be used. 
         [0019]    Band, as used below, may mean the inner band  120 , the outer band  122  or each of the inner band  120  and outer band  122 . An enclosure  142  may be associated with the non-flowpath side  126 ,  130  of the band. The enclosure  142  may be formed integrally with the band or may be attached in any manner known in the art, such as, but not limited to, brazing. In one exemplary embodiment, the enclosure  142  may extend between the non-flowpath side  126 ,  130  of the band and the flange  132 . The enclosure  142 , non-flowpath side  126 ,  130  of the band and the flange  132  may define a plenum  144 . The plenum  144  may receive cooling air through one or more holes  146  in the flange  132 . Air may be routed to the one or more holes  146  through any manner known in the art. As shown in  FIG. 6 , the plenum  144  may be formed as two cavities  148  and  150  in the non-flowpath side  126 ,  130  of the band with a channel  152  connecting therebetween. 
         [0020]    In one exemplary embodiment, the band may have a plurality of cooling holes  154  formed therein. The plurality of cooling holes  154  may have an inlet  156  at the plenum  144  for receiving cooling air therefrom. The plurality of cooling holes  154  may have an outlet  158  for providing film cooling to the flowpath side  124 ,  128  of the band. In one exemplary embodiment, as shown in  FIG. 4 , the outlets  158  of the plurality of cooling holes  154  may be flared to increase the amount of film cooling. A plurality of cooling holes  160  may be formed in the enclosure  142 . The plurality of cooling holes  160  may have an inlet  162  at the plenum  144  for receiving cooling air therefrom. The plurality of cooling holes  160  may have an outlet  164  for providing impingement cooling to the non-flowpath side  126 ,  130  of the band. In one exemplary embodiment, a discourager  166  may be associated with the enclosure  142 . The discourager  166  may be formed integrally with said enclosure  142  or may be attached in any manner known in the art, such as, but not limited to, brazing. The discourager  166  may act as a shield for the non-flowpath side  126 ,  130  of the band to keep hot gases away therefrom. In one exemplary embodiment, the enclosure  142 , plenum  144 , and associated cooling holes may be located on the aft side of the band. 
         [0021]    By providing cooling in these areas, the metal temperature may be reduced, leading to less distress and less likelihood of forming a crack or hole. As such, the turbine nozzle segment will last longer leading to less repairs and/or replacements over time for the gas turbine engine. 
         [0022]      FIG. 8  illustrates an exemplary embodiment of a repair procedure. An engine-run turbine nozzle segment may be provided at step  300 . The turbine nozzle segment may be coated with a thermal barrier coating. The turbine nozzle segment may be cleaned at step  302 . Cleaning may include grit blasting that may remove any corrosion from engine use. Once the turbine nozzle segment is cleaned, the coating may be removed at step  304 . This step may be skipped should the nozzle segment not have a coating. An acid bath may be used to strip the coating. Any acid known in the art may be used. Once the coating is removed, the nozzle segment may be inspected at step  306  to look for any cracks or distressed areas in the base metal. If cracks are found, the damaged material may be removed at step  308 . In one exemplary embodiment, a damaged area may be the enclosure  142  and/or the discourager  166 . The enclosure  142  and discourager  166  may be removed together by placing the nozzle segment into an oven and heating the nozzle segment to a temperature high enough to melt the braze material used to attach the enclosure  142  and discourager  166  to the band. This will allow the enclosure  142  and discourager  166  to be removed without damaging the segment to segment braze material, which has a higher melting temperature than the braze used for the enclosure  142 . In another exemplary embodiment, the discourager  166  may be cut away from the enclosure  142  or the enclosure  142  and discourager may be cut away from the band, depending on which components are damaged. 
         [0023]    At step  310 , a new enclosure  142  and/or discourager  166  may be attached to the nozzle segment. Any other damaged areas of the nozzle segment may be repaired at the same time, either through replacements or through weld repairs where material is added to a damage area and then formed to normal size. The component may be inspected again at step  312 . Cooling holes  160  may be formed integrally with the enclosure. Alternatively, at step  314 , the cooling holes  160  may need to be drilled if only the discourager  166  was replaced in step  310 . The holes  160  may be formed in any manner known in the art, such as, but not limited to, electrodischarge machining, electrochemical machining, laser drilling, mechanical drilling, or any other similar manner. Once all repairs are complete, a new thermal barrier coating may be formed at step  316 . 
         [0024]    This written description discloses exemplary embodiments, including the best mode, to enable any person skilled in the art to make and use the exemplary embodiments. The patentable scope is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.