Abstract:
A shroud for surrounding a portion of a turbine flow path having improved cooling and durability is disclosed. The shroud includes a plurality of generally axial cooling holes spaced a substantially equal distance apart and a plurality of generally circumferential cooling holes oriented generally perpendicular to the generally axial cooling holes. The generally circumferentially cooling holes are spaced a non-uniform distance apart so as to provide cooling to selected portions of shroud sidewalls to lower shroud operating temperatures and improve shroud durability.

Description:
BACKGROUND OF THE INVENTION  
   This invention generally relates to gas turbine engines and more specifically to a shroud section that surrounds a stage of rotating airfoils in the turbine of a gas turbine engine. 
   A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mix the compressed heated air with fuel and contain the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which drives the compressor, before exiting the engine. A portion of the compressed air from the compressor bypasses the combustors and is used to cool the turbine blades and vanes that are continuously exposed to the hot gases of the combustors. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity. 
   In the turbine section of the engine, alternating stages of rotating and stationary airfoils are present through which the hot combustion gases expand as they turn the rotating stages of the turbine. In order to maximize the performance of the turbine, it is critical to maximize the amount of hot combustion gases passing through the airfoils, and not leaking around the airfoils, nor being used to cool the airfoils. To prevent leakage around stages of rotating airfoils, or turbine blades, shroud segments are used that conform to the radial profile of the turbine stage and are sized such that when the blade is rotating and at its operating temperature, the gap between the turbine blade tip and the shroud segment is minimized. 
   Given that operating temperatures within the turbine typically exceed 2000 degrees F. it is necessary to provide a source of cooling to the blades, vanes, and shroud segments adjacent the rotating blades so that these components are maintained within their material operating limits. Of particular concern with respect to the present invention is cooling of the shroud segments that encompass the rotating turbine blades. However, while it is necessary to cool the shroud segments, any air directed to cool the shroud segments does not pass through the turbine, thereby reducing the turbine efficiency. It is imperative that this cooling air, which is typically drawn from the engine compressor, be a minimal amount and used most effectively to cool as much of the exposed shroud surface as possible. An example of a shroud segment for a gas turbine engine employing a form of cooling of the prior art is shown in perspective view in  FIG. 1 . Shroud  10  includes an inner surface  11  that faces directly towards the tips of the rotating turbine blades (not shown) and an outer surface  12  in spaced relation to inner surface  11 . Extending axially through the shroud thickness between inner surface  11  and outer surface  12  and exiting from shroud aft face  13  is a plurality of cooling holes  14 . A cooling fluid, such as compressed air, enters cooling holes  14  from air inlets  15  and cools the shroud  10  as it passes through cooling holes  14 . In this configuration, the edges  16  and  17  of shroud  10  do not receive any dedicated cooling. Shrouds are typically segmented, creating edges  16  and  17 , in order to allow for differing thermal expansion between shroud  10  and the engine case in which the shrouds are mounted. Inspection of prior art shrouds having this cooling configuration indicate excessive heat load along edges  16  and  17 , especially along the axial region of shroud  10  where the turbine blade is located. 
   In order to overcome the shortfalls of the prior art shroud design, it is necessary to provide a shroud for a gas turbine engine which addresses the heat load issues found in the prior art design, including providing sufficient cooling to the edges of the turbine shroud. Providing sufficient cooling to the edge regions where it is most needed will ensure that the heat load is reduced in the effected areas thereby extending the life of turbine shroud segments. 
   SUMMARY OF THE INVENTION  
   The present invention provides an improved shroud that is designed to surround a portion of a turbine. The shroud comprises first and second contoured surfaces, forward and aft faces, and first and second sidewalls. The shroud also comprises a plurality of generally axial cooling holes extending through the shroud thickness and a plurality of generally circumferential cooling holes oriented generally perpendicular to the axial cooling holes. The generally circumferential cooling holes are spaced a non-uniform distance apart so as to provide cooling to selected portions of first and second sidewalls. For the preferred embodiment generally circumferential cooling holes are concentrated higher proximate the axial position of the turbine blade, which imparts the highest heat load to the shroud. The generally axial cooling holes receive their cooling fluid preferably from a plurality of first feed holes, with each feed hole supplying the cooling fluid to an individual generally axial cooling hole. As for the plurality of generally circumferential cooling holes, they receive the cooling fluid preferably from a plurality of openings where each opening directs cooling fluid to multiple circumferential holes. It is preferred that the cooling fluid is air. However, other fluids may be used if available and desirable. 
   The present invention overcomes the shortfalls of the prior art by providing a shroud configuration that provides enhanced and dedicated cooling to previously un-cooled regions of the turbine shroud, specifically the shroud sidewalls. Furthermore, the circumferential cooling holes are spaced such that additional cooling air is directed to the highest temperature regions of the shroud in order to maximize the cooling efficiency. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS  
       FIG. 1  is a perspective view of a turbine shroud of the prior art. 
       FIG. 2  is a perspective view of a turbine shroud in accordance with the preferred embodiment of the present invention. 
       FIG. 3  is a section view of a turbine shroud in accordance with the preferred embodiment of the present invention. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
   The preferred embodiment will now be described in detail with specific reference to  FIGS. 2 and 3 . A shroud  20  for surrounding a portion of a gas turbine engine flow path is shown in perspective view in  FIG. 2  and in a section view in  FIG. 3 . Shroud  20  comprises a number of features including a first surface  21  having a first contour and a second surface  22  having a second contour with second surface  22  located radially outward of first surface  21  thereby establishing thickness  23  therebetween. First contour and second contour are defined by the diameter of the turbine enclosed by shrouds  20 , and will therefore vary in size by design. Shroud  20  further comprises forward face  24  and aft face  25 , which are spaced in axial relation and extend radially between first surface  21  and second surface  22 . Extending generally axially between forward face  24  and aft face  25  and spaced in circumferential relation are first sidewall  26  and second sidewall  27 . An additional feature of shroud  20  is a first row of hooks  28  that extend radially outward from second surface  22  proximate forward face  24 . A plurality of hooks is used in order to secure the shroud to an engine casing that surrounds the turbine section. Typically for structural integrity, hooks  28  are formed integral with shroud  20 . It is common practice in the gas turbine industry to investment cast shrouds  20 , including hooks  28 , and then machine in other features of shroud  20 . One such feature typically machined into a cast shroud is plurality of generally axial cooling holes  29 , which for shroud  20  extend generally axially through the shroud from proximate first row of hooks  28  to aft face  25  and are preferably spaced a substantially equal distance apart. 
   An improvement of the present invention to shroud  20  is a plurality of generally circumferential cooling holes  30  that are oriented generally perpendicular to plurality of generally axial cooling holes  29 . Plurality of generally circumferential cooling holes  30  are spaced a non-uniform distance apart to provide dedicated cooling to regions of first sidewall  26  and second sidewall  27 . An especially high heat load is subjected to shroud  20  proximate first sidewall  26  compared to that of second sidewall  27 . This is due to the direction from which the upstream turbine vanes direct the hot combustion gases onto the turbine blades within shrouds  20 . For this particular shroud design, hot gases are directed from upstream turbine vanes at angle from the forward face  24  and first sidewall  26  towards the aft face  25  and second sidewall  27  (see arrows in  FIG. 2  for flow direction). As a result more cooling holes  30  are required along first sidewall  26  than second sidewall  27 . For this particular shroud, twice as many cooling holes  30  of equal diameter are required for first sidewall  26 . As one skilled in the art of turbine cooling will understand, the exact quantity and size of cooling holes  30  are a function of the cooling required, available cooling air, and operating conditions. A cooling fluid, preferably compressed air, flows through generally axial cooling holes  29  and generally circumferential cooling holes  30 . The cooling fluid is directed to generally axial cooling holes  29  by a plurality of first feed holes  31  in second surface  22 . 
   An additional feature of shroud  20  is plurality of openings  32  located in second surface  22 . Each of plurality of openings  32  has an axial length and a circumferential width with the axial length being greater than the circumferential width. Openings  32  are sized such that each opening is in fluid communication with multiple circumferential cooling holes  30 . The quantity of openings  32  can vary depending on the size of shroud  20  and the quantity of circumferential cooling holes  30  that are fed a cooling fluid from opening  32 . For the preferred embodiment disclosed in the present invention, three openings proximate both first sidewall  26  and second sidewall  27  are utilized. Depending on the size of openings  32  and shroud geometry, openings  32  can be cast into shroud  20  or machined into shroud  20  while machining other features such as cooling holes  29  and  30 . It is preferred that openings  32  are sized with the disclosed axial length and circumferential width relationship for cost and structural reasons. Specifically, it is more cost effective to machine slots into second surface  22  than to drill individual feed holes for directing cooling fluid to each of plurality of circumferential cooling holes  30 . Furthermore, due to the close proximity of plurality of circumferential cooling holes  30 , placing an individual feed hole for each circumferential cooling hole would introduce areas of high stress concentrations at the interface of the circumferential cooling hole and individual feed hole. 
   A further feature of shroud  20  in accordance with the preferred embodiment is a second row of hooks  33  that extend radially outward from second surface  22  proximate aft face  25 . Both second row of hooks  33  and first row of hooks  28  preferably comprises three hooks as shown in  FIG. 2 . Hooks  28  and  33  are designed and spaced such that shroud  20  is held in place within the gas turbine engine by hooks  28  and  33 . 
   The present invention as disclosed herein provides a turbine shroud geometry with improved cooling to regions of the shroud previously uncooled or inadequately cooled. Adequate cooling is especially important along regions of the shroud exposed to the high heat load created by passing rotating turbine blades. 
   While the invention has been described in what is known as presently the 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 within the scope of the following claims.