Patent Abstract:
A turbine blade is disclosed having a tip shroud that includes internal passages through which cooling air is flowed to minimize creep. The cooling air is provided to the shroud through dedicated cooling passageways which include tube inserts that restrict the transfer of heat from the airfoil portion of the turbine blade to the cooling air within the tube as the cooling air passes through the airfoil portion.

Full Description:
BACKGROUND OF THE INVENTION  
         [0001]    The present invention relates to a blade for a gas turbine, and more specifically, to the cooling of a gas turbine blade shroud.  
           [0002]    A gas turbine is typically comprised of a compressor section, a combustor section and a turbine section. The compressor section produces compressed air. Then fuel is mixed with some of the compressed air and burned in the combustor section. The compressed, high temperature gas produced in the combustor section is then expanded through rows of stationary vanes and rotating blades in the turbine section to produce power in the form of a rotating shaft.  
           [0003]    Each of the rotating blades has an airfoil portion and a root portion that connects it to a rotor. Since the blades are exposed to the compressed, hot gas discharging from the combustor section, the turbine blades must be cooled to prevent failure. Usually this cooling is done by taking a portion of the compressed air produced by the compressor and using it as cooling air in the turbine section to cool turbine blades. The cooling air enters each cooled turbine blade through its root, and flows through radial passageways in the airfoil portion of the blades. While in many cooled turbine blades, the radial passageways discharge the cooling air radially outward at the blade tip, some turbine blades incorporate shrouds that project outwardly from the airfoil at the blade tip. These shrouds prevent hot gas leakage past the blade tips, and may also be used to dampen blade vibration that tends to occur during normal operation of gas turbine engines. Unfortunately, excessive creep and creep failures can occur in blade shrouds due to the high operating temperatures.  
           [0004]    While the known methods of cooling turbine blades are generally successful at cooling the airfoil portions of turbine blades, designs for cooling shrouds have produced mixed results. In some designs, cooling air discharged from the radial passages at the blade tip flows over the radially outward facing surface of the shroud. Although this provides some cooling, it is often insufficient to adequately cool the shroud due to heating of the cooling air in the airfoil passageways.  
           [0005]    Another design includes incorporating cooling passages into each shroud, with the cooling passages extending approximately parallel to the radially inward facing surface of the shroud. These passages, which connect to one or more of the radial passageways, divert cooling air from the airfoil passageways so that it flows through the cooling passages in the shroud, thereby lowering the operating temperature of the shroud. While this method of internally cooling the shroud is generally more effective than flowing cooling air over the radially outward facing surface of the shroud, the heat transfer rate from the shroud to the cooling air in the passages may be insufficient to prevent excessive creep at certain operating conditions.  
           [0006]    What is needed is a turbine blade having a shroud that is sufficiently cooled to prevent excessive creep at all engine operating conditions.  
         SUMMARY AND OBJECTS OF THE INVENTION  
         [0007]    It is therefore an object of the present invention to provide a turbine blade having a shroud that is sufficiently cooled at all engine operating conditions to prevent the excessive creep that can occur in turbine shrouds when turbine blades are exposed to high stress and very high operating temperatures.  
           [0008]    According to the preferred embodiment of the present invention, a turbine blade is disclosed having a root portion with a cooling fluid cavity therein, a platform connected to the root portion, an airfoil portion extending from the platform, the airfoil portion includes at least one cooling passageway extending substantially radially through the airfoil, and at least one cooling hole extending substantially radially through the airfoil, with the one cooling passageway and the cooling hole each defined by an inner wall having an inlet for receiving a flow of cooling fluid from the cavity. The turbine blade further includes a shroud projecting outwardly from the airfoil and has a radially inward facing surface, a radially outward facing surface, and a shroud edge extending therebetween, at least one cooling fluid outlet adjacent the edge, and at least one cooling passage between the radially inward facing surface and the radially outward facing surface. The cooling passage is approximately parallel to the radially inward facing surface, and a tube is located within the cooling hole. The tube has an outer wall, a first end adjacent the inlet and a second end radially outward therefrom. The cooling passage communicates with the inlet through the tube, and standoff means between the inner wall of the cooling passageway and the outer wall of the tube maintain the inner wall of said cooling passageway in spaced relation to said outer wall of the tube to minimize heat transfer between the airfoil and the tube.  
           [0009]    The above, and other objects, features and advantages of the present invention will become apparent from the following description read in conjunction with the accompanying drawings. 
       
    
    
     BRIEF DESCRIPTION OF DRAWINGS  
       [0010]    [0010]FIG. 1 shows a turbine blade of the present invention, with certain features shown in phantom lines.  
         [0011]    [0011]FIG. 2 shows a cross-sectional view of the airfoil portion of the present invention taken along line A-A of FIG. 1.  
         [0012]    [0012]FIG. 3 shows a cross-sectional view of a cooling passageway and tube taken along line B-B of FIG. 2.  
         [0013]    [0013]FIG. 4 is a plan view of the shroud of the present invention showing the cooling passageways, cooling passages, and cooling fluid outlets.  
         [0014]    [0014]FIG. 5 shows a cross-sectional view of the shroud of the present invention taken along line C-C of FIG. 4.  
         [0015]    [0015]FIG. 6 is a cross-sectional view similar to FIG. 3, showing a first alternate embodiment of the present invention.  
         [0016]    [0016]FIG. 7 is a cross-sectional view similar to FIG. 3, showing a second alternate embodiment of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0017]    The present invention is relates to cooled turbine blades of the type used in gas turbine engines in which cooling air is supplied by the compressor of the gas turbine and is directed into the root of the cooled turbine blades through the rotors. These methods of getting the compressed air to the turbine blade roots will not be addressed in this description since these methods are well known in the art.  
         [0018]    As shown in FIG. 1, the turbine blade  10  of the present invention includes a root portion  12  having a cooling fluid cavity  14  therein. A platform  16  is connected to the root portion, and an airfoil portion  18  extends away from from the platform  16  in a direction that is substantially parallel to a first radial direction  20 . The airfoil portion  18  includes at least one, and preferably a plurality of cooling passageways  22  extending substantially radially through the airfoil portion  18 . Each cooling passageway  22  has an inlet  24  for receiving a flow of cooling fluid from the cavity  14 . In addition to the cooling passageways  22 , the airfoil  18  preferably includes cooling holes  26  extending substantially radially through the airfoil portion  18 . Each cooling hole  26  also has an inlet  28  for receiving a flow of cooling fluid from the cavity  14 . A shroud  30  extends outwardly from the airfoil  18  adjacent the end of the airfoil  18  opposite the platform  16 .  
         [0019]    As shown in FIG. 2, a tube  32  is located within each cooling passageway  22 . By contrast, the cooling holes  26  do not contain insulating tubes, since this would necessarily impair their ability to cool the airfoil portion  18  of the turbine blade  10 . Each tube  32  has an outer wall  34  and an internal wall  36 .  
         [0020]    Referring now to FIG. 3, each insulating tube  32  has a first end  38  adjacent the inlet  24  of the passageway  22  in which it is located. In the preferred embodiment, standoff means extend from the inner wall  42  of the cooling passageway  22 . The standoff means comprise at least one, and preferably a plurality of, protrusions  40  extending inwardly from the inner wall  42  of of the passageway  22 . Each protrusion  40  may be annular and therefore entirely encircle the tube  32 , or each protrusion  40  may be nearly a localized “bump”, which cooperates with other the other protrusions to maintain the relative position of the tube  32  in the cooling passageway  22 . Each protrusion  40  contacts the outer wall  34  of the tube  32 , thereby maintaining the inner wall  42  of the cooling passageway  22  in spaced relation to the outer wall  34  of the insulating tube  32 . As those skilled in the art will readily appreciate, minimizing the contact area between the tube  32  and the inner wall  42  minimizes heat transfer between the airfoil portion  18  and the insulating tube  32 .  
         [0021]    As shown in FIG. 4, the shroud  30  preferably has a “Z-notch” configuration of the type known in the art. Each shroud  30  includes at least one, and preferably a plurality of cooling passages  44 . Each cooling passage  44  has a cooling fluid outlet  46  adjacent an edge  48  that forms a portion of the Z-notch. Each cooling passage  44  communicates with an inlet  24  through one of the tubes  32 . As shown in FIG. 5, each shroud  30  has a radially inward facing surface  50 , a radially outward facing surface  52 , and a shroud edge  48  extending therebetween. Each cooling passage  44  is located between the radially inward facing surface  50  and the radially outward facing surface  52 . The cooling passages  44  are approximately parallel to the radially inward facing surface  50 .  
         [0022]    Each tube  32  has a second end  54  radially outward from the first end  38  thereof. The second end  54  abuts a tube retention plug  56 . The tube retention plug  56  has an internal flowpath  58 , including a flowpath inlet  59  and at least one flowpath outlet  60 . The second end  54  of the tube  32  is preferably sealingly fixed to the tube retention plug  56  at the flowpath inlet  59 . Each cooling passage  44  is in fluid communication with one of the tubes  32  through the internal flowpath  58  of one of a tube retention plug  56 . The internal flowpath preferably includes metering means  62  for restricting fluid flow from the tube  32  to each cooling passage  44 .  
         [0023]    As shown in FIG. 4, the preferred embodiment of the present invention has at least two cooling passageways  22  and a plurality of cooling passages  44 . Although the cooling fluid outlet  46  is shown in in the radially outward facing surface  52  of FIG. 5, it is to be understood that the cooling fluid outlet  46  may be located in the shroud edge  48  if it is desirable to flow cooling fluid into the gap  64  between the shrouds of adjacent turbine blades  10 . Likewise, if film cooling is desired along the edge  48  at the radially inward facing surface  50 , the cooling fluid outlet  46  may be located in the radially inward facing surface  50  immediately adjacent the edge  48 .  
         [0024]    [0024]FIG. 6 shows a first alternate embodiment of the present invention, which is similar to the design of the preferred embodiment, except that the standoff means are different and a flange may be added to the cooling tube  32 . In the first alternate embodiment, the inner wall  42  of the cooling passageway  22  is smooth, and at least one, and preferably a plurality of, protrusions  66  extend from the tube  32  and contact the inner wall  42  of the cooling passageway  22 . As those skilled in the art will readily appreciate, the protrusions  66  maintain that tube  32  in spaced relation to the inner wall  42  of the cooling passageway  22 , thereby minimizing heat transfer between the airfoil portion  18  and the tube  32 . If the protrusions  66  are not annular, cooling air may be able to pass between the inner wall  42  of the cooling passageway  22  and the tube  32 . Therefore, in the first alternate environment, it is preferable to provide an annular flange  68  at the inlet  24  to the cooling passageway  22  to direct the cooling air into the tube  32 , and prevent cooling air from flowing between the inner wall  42  of the cooling passageway  22  and the tube  32 .  
         [0025]    [0025]FIG. 7 shows a second alternate embodiment of the present invention, which likewise is similar to the design of the preferred embodiment except for the standoff means and the cooling tube flange. As in the first alternate embodiment, the inner wall  42  of the cooling passageway  22  is smooth, and at least one, and preferably a plurality of, protrusions  70  extend from the tube  32  and contact the inner wall  42  of the cooling passageway  22 . In the second alternate embodiment, the protrusions  70  are preferably annular, so that each protrusion  70  acts to prevent the flow cooling air through the between the inner wall  42  of the cooling passageway  22  and the tube  32 . The second alternate embodiment also preferably includes a flange  72  that performs the same functions as the flange  68  in the first alternate embodiment. However, since each protrusion  70  in the second alternate embodiment impedes the flow of cooling air between the inner wall  42  of passageway  22  and the tube  32 , flange  72  is not as critical to the overall performance of the present invention. In fact, the flange  72  may be identical to the protrusions  70 .  
         [0026]    Although the preferred embodiments of the present invention have been described with reference to the accompanying drawings, it is to be understood that the invention is not limited to those precise embodiments, and that various changes and modifications may be effected therein by one skilled in the art without departing from the scope or spirit of the invention as defined in the appended claims.

Technology Classification (CPC): 5