Abstract:
A segment of a component for use in a gas turbine includes a leading edge; a trailing edge; a pair of opposed lateral sides between the leading and trailing edges; and a seal slot provided in each lateral side. The seal slot includes a surface having a channel extending in an axial direction defined from the leading edge to the trailing edge, at least one inlet to the channel, and at least one outlet from the channel. The at least one outlet is spaced downstream from the at least one inlet in the axial direction. The segment may be an inner shroud segment or a nozzle segment.

Description:
The present invention relates to shrouds and nozzles for gas turbines and, more particularly, to arrangements for cooling shrouds and nozzles of gas turbines. 
     BACKGROUND OF THE INVENTION 
     Shrouds employed in a gas turbine surround and in part define the hot gas path through the turbine. Shrouds are typically characterized by a plurality of circumferentially extending shroud segments arranged about the hot gas path, with each segment including discrete inner and outer shroud bodies. Conventionally, there are two or three inner shroud segments for each outer shroud segment, with the outer shroud segments being secured to the stationary inner shell or casing of the turbine and the inner shroud segments being secured to the outer shroud segments. The inner shroud segments directly surround the rotating parts of the turbine, i.e., the rotor wheels carrying rows of buckets or blades. 
     Because the inner shroud segments are exposed to hot combustion gases in the hot gas path, systems for cooling the inner shroud segments are oftentimes necessary to reduce the temperature of the segments. This is especially true for inner shroud segments in the first and second stages of a turbine that are exposed to very high temperatures of the combustion gases due to their close proximity to the turbine combustors. Heat transfer coefficients are also very high due to rotation of the turbine buckets or blades. 
     To cool the shrouds, typically relatively cold air from the turbine compressor is supplied via convection cooling holes that extend through the segments and into the gaps between the segments to cool the sides of the segments and to prevent hot path gas ingestion into the gaps. The area that is purged and cooled by a single cooling hole is small, however, because the velocity of the cooling air exiting the cooling hole is high and the cooling air diffuses into the hot gas flow path. 
     Typically, the post-impingement air leaks into the gas path between two inner shrouds, through hard/cloth seals located on the seal slot surface. Shroud slash faces, in particular, above the bucket region, are the life-limiting regions, mainly due to oxidation. This is caused by the continuous ingestion of hot gases thrown by the bucket towards the shroud inter-segment gaps. Traditional cooling methods use cooling holes along the slash face drilled from post-impingement cold section, or discrete perpendicular channels machined along the length of the seal slot, which improves the slash face cooling to certain extent, but whose effects are very localized as they do not cover the entire length of low-life slash face region. 
     Another component of gas turbines that includes seal slots are nozzles. A nozzle may be formed by a plurality of sections, or segments, and seals between adjacent segments. Service run nozzles in a gas turbine may have distorted sidewalls as a result of previous weld repairs or due to stress relief during service. Creep strain due to applied loads at operating temperatures may also contribute to distortion. This movement of the sidewalls will cause the seal slots that are contained within the sidewalls to be out of position relative to engine center. 
     If the sidewalls are not pressed back into position, the seal slots between adjacent segments would not be aligned with each other, and it may prove impossible to fit the seals in place. Alternatively, it may be possible to force the seals into the slots but this would lock the nozzle segments together such that they could not move or “float” relative to each other. This float is necessary to allow for thermal expansion and to ensure the segments load up against the sealing faces (hook fit and chordal hinge) during operation. If they are locked together, it is likely they will be skewed and will not load against their sealing faces. This will result in compressor discharge air leaking directly into the hot gas path and will reduce the amount of air available for combustion and for cooling of the nozzle. The result of reduced air for combustion will be lower performance of the turbine and increased emissions. A reduction in available cooling air will result in increased oxidation of the nozzle due to a resultant higher metal temperature and the lack of cooling will also cause changes to thermal gradients within the nozzle leading to increased cracking of the part. This will increase subsequent repair costs and may reduce the life of the parts. 
     Misaligned sidewalls may also result in flow path steps. The hot gas will not have a smooth path but will be tripped by the mismatch between adjacent nozzle segments, resulting in turbulent flow and reduced energy of the gas stream, thereby reducing performance. Turbulent flow also increases thermal transfer to the nozzle and so will raise the metal temperature, leading to increased oxidation and cracking. 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to one embodiment, a segment of a component for use in a gas turbine comprises a leading edge; a trailing edge; a pair of opposed lateral sides between the leading and trailing edges; and a seal slot provided in each lateral side. The seal slot comprises a surface having a channel extending in an axial direction defined from the leading edge to the trailing edge, at least one inlet to the channel, and at least one outlet from the channel, wherein the at least one outlet is spaced downstream from the at least one inlet in the axial direction. 
     According to another embodiment, a gas turbine comprises at least one of an inner shroud and a nozzle, wherein at least one of the inner shroud and the nozzle comprises a plurality of circumferentially arranged segments, and each segment comprises a leading edge, a trailing edge, a pair of opposed lateral sides between the leading and trailing edges; and a seal slot provided in each lateral side, the seal slot comprising a surface, the surface comprising a channel extending in an axial direction defined from the leading edge to the trailing edge, at least one inlet to the channel, and at least one outlet from the channel, wherein the at least one outlet is spaced downstream from the at least one inlet in the axial direction. 
     According to yet another embodiment, a method of cooling a component of a gas turbine is provided. The component comprises a plurality of segments circumferentially arranged. Each segment comprises a leading edge, a trailing edge, a pair of opposed lateral sides between the leading and trailing edges, and a seal slot provided in each lateral side. The component further comprises a seal on each seal slot. The method comprises directing cooling air that leaks into the seal slot below the seal through at least one inlet into a channel formed in a surface of the seal slot, wherein the channel extends in an axial direction defined from the leading edge to the trailing edge; directing the leaking cooling air along the channel; and directing the leaking cooling air out of the channel through at least one outlet, wherein the at least one outlet is spaced downstream from the at least one inlet in the axial direction. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a front perspective view of an inner shroud segment; 
         FIG. 2  is a rear perspective of the inner shroud segment of  FIG. 1 ; 
         FIG. 3  is a side perspective of the inner shroud segment of  FIGS. 1 and 2 ; 
         FIG. 4  is a side perspective of another inner shroud segment; 
         FIG. 5  is a perspective view of a gas turbine nozzle section; 
         FIG. 6  is a plan view of a seal slot surface according to an embodiment of the invention; 
         FIG. 7  is a plan view of a seal slot surface according to another embodiment of the invention; and 
         FIG. 8  is a plan view of a seal slot surface according to a further embodiment of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIGS. 1-3 , an inner shroud segment  2  comprises a leading edge  4  and a trailing edge  6 . The inner shroud segment  2  is configured to be connected to an outer shroud segment by a leading edge hook  8  and a trailing edge hook  10 . 
     The inner shroud segment  2  comprises impingement cavities, or plenums,  12  which receive relatively cold air from the turbine compressor to cool the inner shroud segments. As shown in  FIG. 1 , trailing edge convection cooling apertures  14  extend through the inner shroud segment  2 , and as shown in  FIG. 2 , leading edge convection cooling apertures  16  are provided adjacent the leading edge  4 . 
     Referring still to  FIGS. 1-3 , the inner shroud segment  2  may comprise a seal slot  18  configured to receive a hard/cloth seal located on the seal slot surface  22 . Typically, the post-impingement air leaks into the gas path between two inner shroud segments and through the hard/cloth seals located on the seal slot surface  22 . The post-impingement leakage/cooling air enters the seal slot  18  below the hard/cloth seals on the seal slots  18  and exits into the hot gas path, thus providing active cooling closer to the slash faces  20  of the inner shroud segments. The slash faces  20  are provided on opposed lateral sides of the inner shroud segment  2 . 
     Referring to  FIG. 4 , discrete channels  24  are provided in the seal slot surface  22 . The post-impingement leakage/cooling air enters perpendicular inlet channels  24  below the hard/cloth seals on the seal slots  18  and provides active cooling to the slash face  20 . As used herein, the term perpendicular refers to a direction perpendicular to the axial direction of the inner shroud segment defined from the leading edge to the trailing edge in a direction from an upstream position to a downstream position of a hot gas path through the turbine shroud. The cooling provided by the inlet channels  24  is localized and does not cover the entire length of the slash face region. 
     Referring to  FIG. 5 , a section or segment of a gas turbine nozzle includes an outer wall  42 , an inner wall  46 , and an airfoil  44  between the walls  42 ,  46 . The nozzle segment includes a leading edge  4  and a trailing edge  6 . The section also includes a number of seal slots  18  provided in opposed lateral sides of the nozzle segment. The seal slots  18  retain the end face seals (sometimes referred to as spline seals or slash face seals) that seal between adjacent nozzle segments and prevent the compressor discharge air leaking into the hot gas path and prevent ingestion of hot gas into the component. 
     Referring to  FIG. 6 , according to an embodiment of the invention, the seal slot surface  22  comprises a plurality of perpendicular inlet channels  28 . The post-impingement leakage/cooling air  26  enters the multiple perpendicular inlet channels  28  and then flows axially in a channel  30 , and then enters perpendicular exit channels  32  into the hot gas path  34 . As used herein, the term axial refers to the direction of the inner shroud segment from the leading edge to the trailing edge in a direction from an upstream position to a downstream position of the hot gas path through the turbine. 
     As shown in  FIG. 6 , the exit channels  32  are located alternately from the inlet channels  28 . This configuration reduces the possibility that combustion gases from the hot gas path  34  may enter the seal slot of the inner shroud segment. It should be appreciated, however, that the inlet channels  28  and the exit channels  32  may be coaxial to each other. It should also be appreciated that the inlet channels  28  and/or the outlet channels  32  may not be perpendicular to the axial channel  30 , but may instead be provided at an angle to the axial channel  30 . It should be further appreciated that the number of inlet channels may be different from the number of outlet channels, or that the widths and/or lengths of the inlet channels and/or the outlet channels may be different from each other. 
     Referring to  FIG. 7 , a seal slot surface  22  according to another embodiment comprises a plurality of perpendicular inlet channels  28 . The post-impingement leakage/cooling air  26  enters the inlet channels  28  and flows into the channel  30  and then flows out the perpendicular exit channels  32  into the hot gas path  34 . As shown in  FIG. 7 , the exit channels  32  are provided after the inlet channels  28  in the axial direction of the seal slot surface  22 . This configuration provides robust cooling in cases where the leading edge backflow margin is low because it prevents hot gases from short-circuiting through the exit channels  32  near the leading edge of the segment. 
     Referring to  FIG. 8 , a seal slot surface  22  according to another embodiment includes a channel  36 . The leakage/cooling air  26  enters the channel at inlet  38  and exits the channel  36  at outlet  40 . The channel  36  may take a zig-zag configuration in the seal slot surface  22 . Alternatively to, or in combination with, the zig-zag configuration, the channel may include a serpentine configuration Although each portion, or segment, of the channel  36  is shown as linear in  FIG. 8 , it should be appreciated that the portions, or segments, may be curved, or curvilinear. The configuration of  FIG. 8  provides an increased convection path length compared to the embodiments shown in  FIGS. 6 and 7 . 
     The channels  30 ,  36  shown in the embodiments of  FIGS. 6-8  provide continuous convective cooling of the seal slot surface  22  closer to the hot surface of the slash face. By providing continuous partial or full length axial convective cooling, the heat transfer coefficient of the post-impingement leakage/cooling air is increased and effective cooling closer to the hot slash face can be achieved. Continuous partial or full length axial convective cooling closer to the hot metal helps to cool the slash face, thus increasing the mechanical life of the inner shroud and/or nozzle segments. As more cooling is provided to the shroud and/or nozzle low life regions, in particular to the slash face length of the shroud segment above the bucket region of the turbine, it is possible to achieve higher mechanical life. 
     The seal slot surfaces of the embodiments shown in  FIGS. 6-8  may be cast with the seal slot of the inner shroud segment or nozzle segment. It should also be appreciated that the embodiments of the seal slot surface  22  shown in  FIGS. 6-8  may be formed by electro-discharge machining of the seal slot surface of an inner shroud or nozzle segment. Existing shroud and/or nozzle segments may thus be modified to include seal slot surfaces having continuous axial channels and an inlet(s) and an outlet(s). 
     The cooling flow along the seal slot channels can be used to cool the slash face metal temperature below certain temperature requirement, resulting in a more uniform metal temperature distribution. By providing continuous partial or full length axial convective cooling, effective cooling closer to the hot slash face can be achieved. The reduction in slash face temperature can increase shroud and nozzle part intervals and achieve higher mechanical life. Since the life-limiting region of the shroud and/or nozzle is targeted, higher mechanical life can be achieved with the increase of HGP intervals. 
     While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.