Patent Abstract:
An open cooling circuit for a gas turbine bucket wherein the bucket has an airfoil portion, and a tip shroud, the cooling circuit including a plurality of radial cooling holes extending through the airfoil portion and communicating with an enlarged internal area within the tip shroud before exiting the tip shroud such that a cooling medium used to cool the airfoil portion is subsequently used to cool the tip shroud.

Full Description:
TECHNICAL FIELD  
         [0001]    This invention relates to a cooling air circuit for a gas turbine bucket tip shroud.  
         BACKGROUND OF THE INVENTION  
         [0002]    Gas turbine buckets have airfoil shaped body portions connected at radially inner ends to root portions and at radially outer ends to tip portions. Some buckets incorporate shrouds at the radially outermost tip, and which cooperate with like shrouds on adjacent buckets to prevent hot gas leakage past the tips and to reduce vibration. The tip shrouds are subject to creep damage, however, due to the combination of high temperature and centrifugally induced bending stresses. In U.S. Pat. No. 5,482,435, there is described a concept for cooling the shroud of a gas turbine bucket, but the cooling design relies on air dedicated to cooling the shroud. Other cooling arrangements for bucket airfoils or fixed nozzle vanes are disclosed in U.S. Pat. Nos. 5,480,281; 5,391,052 and 5,350,277.  
         BRIEF SUMMARY OF THE INVENTION  
         [0003]    This invention utilizes spent cooling air exhausted from the airfoil itself for cooling the associated tip shroud of the bucket. Specifically, the invention seeks to reduce the likelihood of gas turbine tip shroud creep damage while minimizing the cooling flow required for the bucket airfoil and shroud. Thus, the invention proposes the use of air already used for cooling the bucket airfoil, but still at a lower temperature than the gas in the turbine flowpath, for cooing the tip shroud.  
           [0004]    In one exemplary embodiment of the invention, leading and trailing groups of cooling holes extend radially outwardly within the airfoil generally along respective leading and trailing edges of the airfoil. Each group of holes communicates with a respective cavity or plenum in the radially outermost portion of the airfoil. Spent cooling air from the radial cooling passages flows into the pair of plenums and then through holes in the tip shroud and exhausted into the hot gas path. These latter holes can extend within the plane of the tip shroud and open along the peripheral edges of the shroud, or at an angle so as to open through the top surface of the shroud.  
           [0005]    In a second exemplary embodiment, relatively small film cooling holes are drilled through the radial plenum walls on both the pressure and suction side of the airfoil. These holes open on the underside of the shroud, in the area of the shroud fillets. In a variation of this arrangement, the leading and trailing plenums as described above are connected by an internal connector cavity. Preferably, the majority of the cooling holes open along the pressure and suction side in the leading edge area of the blade, with fewer holes opening in the trailing edge area. Covers are joined to the shroud to close the plenums and one or more metering holes are drilled in the respective covers in order to control the cooling air exhaust.  
           [0006]    In a third exemplary embodiment, the individual radial cooling holes within the airfoil are drilled slightly oversize at the tip shroud end. In other words, each cooling hole may be considered to have its own plenum or chamber. Plugs or inserts are joined to the holes to seal the ends of the latter, while shroud cooling holes are drilled directly into the individual plenums and exit either at the top of the shroud or along the underside of the shroud. A metering hole may be required in the various radial cooling hole plugs to insure proper flow distribution.  
           [0007]    In its broader aspects, the invention relates to an open cooling circuit for a gas turbine bucket wherein the bucket has an airfoil portion, and a tip shroud, the cooling circuit comprising a plurality of radial cooling holes extending through the airfoil portion and communicating with an enlarged internal area within the tip shroud before exiting the tip shroud such that a cooling medium used to cool the airfoil portion is subsequently used to cool the tip shroud.  
           [0008]    In another aspect, the invention relates to an open cooling circuit for a gas turbine airfoil and associated tip shroud comprising a plurality of cooling holes internal to the airfoil and extending in a radially outward direction; a first plenum chamber in an outer radial portion of the airfoil, each of the plurality of holes communicating with the plenum; additional cooling holes in the tip shroud, communicating with the plenum, and exiting through the tip shroud.  
           [0009]    In still another aspect, the invention relates to a method of cooling a gas turbine airfoil and associated tip shroud comprising a) providing radial holes in the airfoil and supplying cooling air to the radial holes; b) channeling the cooling air to a plenum in the airfoil; and c) passing the cooling air from the plenum and through the tip shroud.  
           [0010]    Additional objects and advantages of the invention will become apparent from the detailed description which follows.  
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0011]    [0011]FIG. 1 is a partial side section illustrating the turbine section of a land based gas turbine;  
         [0012]    [0012]FIG. 2 is a partial side elevation, in generally schematic form, illustrating groups of radial cooling passages in a turbine blade and tip shroud in accordance with a first exemplary embodiment of the invention;  
         [0013]    [0013]FIG. 3 is a top plan view of a tip shroud in accordance with the first embodiment of the invention;  
         [0014]    [0014]FIG. 4 is a top plan view showing an alternative to the arrangement shown in FIG. 3;  
         [0015]    [0015]FIG. 5 is a top plan view of a turbine airfoil and tip shroud in accordance with a second exemplary embodiment of the invention;  
         [0016]    [0016]FIG. 6 is a section taken along the line A-A of FIG. 5;  
         [0017]    [0017]FIG. 7 is a top plan of an airfoil and tip shroud similar to FIG. 5, but illustrating a connector cavity between the interior plenums;  
         [0018]    [0018]FIG. 8 is a top plan view of a tip shroud in accordance with a third exemplary embodiment of the invention, illustrating shroud cooling holes opening on the top surface of the tip shroud;  
         [0019]    [0019]FIG. 9 is a top plan view of the tip shroud shown in FIG. 8, but illustrating the shroud cooling holes which open along the bottom surface of the tip shroud;  
         [0020]    [0020]FIG. 10 is a section taken along the line  10 - 10  of FIG. 8; and  
         [0021]    [0021]FIG. 11 is a section taken along the line  11 - 11  of FIG. 9. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0022]    With reference to FIG. 1, the turbine section  10  of a gas turbine is partially illustrated. The turbine section  10  of the gas turbine is downstream of the turbine combustor  11  and includes a rotor, generally designated R, with four successive stages comprising turbine wheels  12 ,  14 ,  16  and  18  mounted to and forming part of the rotor shaft assembly for rotation therewith. Each wheel carries a row of buckets B 1 , B 2 , B 3  and B 4 , the blades of which project radially outwardly into the hot combustion gas path of the turbine. The buckets are arranged alternately between fixed nozzles N 1 , N 2 , N 3  and N 4 . Alternatively, between the turbine wheels from forward to aft are spacers  20 ,  22  and  24 , each located radially inwardly of a respective nozzle. It will be appreciated that the wheels and spacers are secured to one another by a plurality of circumferentially spaced axially extending bolts  26  (one shown), as in conventional gas turbine construction.  
         [0023]    Turning now to FIGS. 2 and 3, a turbine bucket includes a blade or airfoil portion  30  and an associated radially outer tip shroud  32 . The airfoil  30  has a first set of internal radially extending cooling holes generally designated  34 , and a second set of five radially extending cooling holes  36 . The first set of cooling holes  34  is located in the forward half of the airfoil, closer to the leading edge  38 , whereas the second set of holes  36  is located toward the rearward or trailing edge  40  of the airfoil. The first set of leading edge cooling holes  34  open to a first cavity or plenum  42  at the radially outermost portion of the airfoil, while trailing edge cooling holes  36  open into a second plenum  44  closer to the trailing edge  40  of the airfoil. The plenums  42  and  44  are shaped to conform generally with the shape of the airfoil, and extend radially into the tip shroud  32 . The plenums are sealed by recessed covers such as those shown at  46 ,  48 , respectively, in FIG. 4. The covers may have metering holes  50 ,  52  for controlling the exhaust rate of the cooling air into the hot gas path.  
         [0024]    In addition, the plenums  42  and  44  can exhaust directly through cooling passages internal to the tip shroud. For example, as shown in FIG. 3, spent cooling air from chamber  42  can exhaust through the edges of the tip shroud via passages  54 ,  56  and  58  which lie in the plane of the shroud  32  and which distribute cooling air within the shroud itself, thus film cooling and convection cooling the shroud. Similarly, plenum  44  communicates with a similar passage  60  in the trailing edge portion of the shroud  32 .  
         [0025]    It will be appreciated that the number and diameter of radial holes in the airfoil will depend on the design requirements and manufacturing process capability. Thus, FIG. 2 shows groups  34 ,  36  of four and three radial holes respectively, whereas FIG. 3 shows both groups to have five radial holes each.  
         [0026]    In FIG. 4, a variation of this embodiment has cooling holes  62 ,  64 ,  66 ,  68 ,  70  and  72  in the tip shroud, in communication with the leading plenum  42 , but angled relative to the plane of the tip shroud so that they exhaust through the top surface  74  of the tip shroud, rather than at the shroud edge. Similarly, cooling holes  76 ,  78  and  80  in communication with the trailing plenum  44  also exhaust through the top surface  74  of the shroud.  
         [0027]    [0027]FIGS. 5 and 6 illustrate a second embodiment of the invention, and, for convenience, reference numerals similar to those used in FIGS. 2 and 3 are used in FIG. 4 where applicable to designate corresponding components, but with the prefix “1” added. Thus, a first set of radially extending internal cooling holes  134  extends radially outwardly through the airfoil, closer to the leading edge  138  of the airfoil, opening at plenum  142 . A similar second set of cooling holes  136  extends radially outwardly within the airfoil, closer to the trailing edge  140  of the airfoil, opening into plenum  144 . A first group of shroud cooling holes  162 ,  164 ,  166  and  168 ,  170 ,  172  and  174  extend from both the pressure and suction sides, respectively, of the plenum  142  to provide film and convection cooling of the underside of the tip shroud  132 , with the cooling holes exiting the airfoil in the area of the tip shroud fillet  82 . A second group of shroud cooling holes  176 ,  178  extend from plenum  144  and open on pressure and suction sides, respectively of the airfoil, again on the underside of the tip shroud. As in the previous embodiment, flow may also be metered out of the plenum covers  146 ,  148  by means of one or more metering holes  150  (FIG. 7). The number of shroud cooling holes exiting on the pressure and suction sides of the shroud may vary as required.  
         [0028]    [0028]FIG. 7 is similar to FIG. 5 but includes a connector cavity  84  extending internally between the leading and trailing plenums  142 ,  144 , respectively. Cooling holes from the plenums exhaust about the tip shroud undersurface as described above. The connector cavity  84  results in most cooling air flowing to the leading edge plenum  142  to exit via cooling holes  162 ,  164 ,  166  and  168 ,  170 ,  172  and  174  arranged primarily along the pressure and suction sides, respectively, of the airfoil in the leading edge region thereof. As in FIG. 6, only two of the cooling holes  176 ,  178  exit in the trailing edge area of the airfoil. This arrangement desirably channels most of the cooling air to the leading edge region of the airfoil, to be washed back across the trailing edge region by the hot combustion gas, thereby providing desirable cooling of the shroud. The metering hole  150  in the cover  146  exhausts all of the spent cooling air which is not otherwise used for direct tip shroud cooling along the undersurface thereof, and dilutes the hot gas flowing over the top of the shroud.  
         [0029]    FIGS.  8 - 11  illustrate a third embodiment of the invention, and, for convenience, reference numerals similar to those used to describe the earlier embodiments are used in FIGS.  8 - 11  where applicable to designate corresponding components, but with the prefix “2” added. A first set of radially extending internal cooling holes  234  extends radially outwardly through the airfoil, closer to the leading edge  238  of the airfoil. A second set of internal cooling holes extends radially outwardly within the airfoil, closer to the trailing edge  240  of the airfoil. Each individual radial cooling hole  234  is drilled or counterbored at its radially outer end to define an individual plenum  242 , while each radial cooling hole  236  is similarly drilled or counterbored to form a similar but smaller plenum  244 . Each enlarged chamber or plenum  242 ,  244  is sealed by a plug or cover  246  (in FIGS. 8 and 9, the plugs or covers  246  are omitted for purposes of clarity). Each plug or cover may be provided with a metering hole  250  to insure proper flow distribution.  
         [0030]    A first group of shroud film cooling holes  262 ,  264 ,  266 ,  268 ,  270 , and  272  extend from the various plenums  242  through the tip shroud and open along the top surface of the tip shroud. Similarly, a second group of film cooling holes  274 ,  276 , and  278  extend from the plenums  244  and also open along the top surface of the tip shroud. Note that film cooling holes  264  and  262  extend from the same plenum, while film cooling holes  270  and  272  extend from the next adjacent plenum. The arrangement may vary, however, depending on particular applications.  
         [0031]    [0031]FIG. 9 illustrates film cooling holes extending from the plenums  242  and  244 , but which open along the underside of the tip shroud, generally along the tip shroud fillet  282 . Thus, film cooling holes  284 ,  286 ,  288 , and  290  extend from two of the plenums  242  and open on the underside of the tip shroud, on both pressure and suction sides of the airfoil. Note that film cooling holes  284  and  290  extend from the same plenum, while a similar arrangement exists with respect to shroud film cooling holes  286  and  288  which extend from the adjacent plenum.  
         [0032]    Shroud film cooling holes  294  and  296  extend from a pair of adjacent plenums  244  associated with radial cooling holes  236  on the opposite side of the tip shroud seal, also along the underside of the tip shroud.  
         [0033]    These arrangements are intended to reduce the likelihood of gas turbine shroud creep damage while minimizing the cooling flow required for the bucket, while more efficiently utilizing spent airfoil cooling air to also cool the tip shroud.  
         [0034]    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.

Technology Classification (CPC): 5