Abstract:
A turbine rotor blade for a gas turbine engine including an airfoil and dovetail for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud, the airfoil comprising: a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge, the pressure sidewall and suction sidewall extending from a root to a tip plate; a pressure tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the pressure tip wall resides approximately adjacent to the termination of the pressure sidewall; a suction tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the suction tip wall resides approximately adjacent to the termination of the suction sidewall; and one or more tip ribs that extend substantially between the pressure tip wall and the suction tip wall.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The present application relates generally to apparatus, methods and/or systems for discouraging cross-flow over turbine airfoil tips. More specifically, but not by way of limitation, the present application relates to apparatus, methods and/or systems related to turbine blade tips that include a squealer tip and/or cross ridges or ribs that discourage cross-flow the blade. 
         [0002]    In a gas turbine engine, it is well known that air is pressurized in a compressor and used to combust a fuel in a combustor to generate a flow of hot combustion gases, whereupon such gases flow downstream through one or more turbines so that energy can be extracted therefrom. In accordance with such a turbine, generally, rows of circumferentially spaced rotor blades extend radially outwardly from a supporting rotor disk. Each blade typically includes a dovetail that permits assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, as well as an airfoil that extends radially outwardly from the dovetail. 
         [0003]    The airfoil has a generally concave pressure side and generally convex suction side extending axially between corresponding leading and trailing edges and radially between a root and a tip. It will be understood that the blade tip is spaced closely to a radially outer turbine shroud for minimizing leakage therebetween of the combustion gases flowing downstream between the turbine blades. Maximum efficiency of the engine is obtained by minimizing the tip clearance or gap such that leakage is prevented, but this strategy is limited somewhat by the different thermal and mechanical expansion and contraction rates between the rotor blades and the turbine shroud and the motivation to avoid an undesirable scenario of having the tip rub against the shroud during operation. 
         [0004]    In addition, because turbine blades are bathed in hot combustion gases, effective cooling is required for ensuring a useful part life. Typically, the blade airfoils are hollow and disposed in flow communication with the compressor so that a portion of pressurized air bled therefrom is received for use in cooling the airfoils. Airfoil cooling is quite sophisticated and may be employed using various forms of internal cooling channels and features, as well as cooling holes through the outer walls of the airfoil for discharging the cooling air. Nevertheless, airfoil tips are particularly difficult to cool since they are located directly adjacent to the turbine shroud and are heated by the hot combustion gases that flow through the tip gap. Accordingly, a portion of the air channeled inside the airfoil of the blade is typically discharged through the tip for the cooling thereof. 
         [0005]    It will be appreciated that conventional blade tip design includes several different geometries and configurations that are meant prevent leakage and increase cooling effectiveness. Exemplary patents include: U.S. Pat. No. 5,261,789 to Butts et al.; U.S. Pat. No. 6,179,556 to Bunker; U.S. Pat. No. 6,190,129 to Mayer et al.; and, U.S. Pat. No. 6,059,530 to Lee. Conventional blade tip designs, however, all have certain shortcomings, including a general failure to adequately reduce leakage and/or allow for efficient tip cooling that minimizes the use of efficiency-robbing compressor bypass air. Improvement in the pressure distribution near the tip region is still sought to further reduce the overall tip leakage flow and thereby increase turbine efficiency. As a result, a turbine blade tip design that alters the pressure distribution near the tip region and otherwise reduces the overall tip leakage flow, thereby increasing the overall efficiency of the turbine engine, would be in great demand. Further, it is also desirable for such a blade tip to enhance the cooling characteristics of the cooling air that is released at the blade tip, as well as, enhancing the overall aerodynamic performance of the turbine blade. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    The present application thus describes a turbine rotor blade for a gas turbine engine including an airfoil and dovetail for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud, the airfoil comprising: a pressure sidewall and a suction sidewall that join together at a leading edge and a trailing edge, the pressure sidewall and suction sidewall extending from a root to a tip plate; a pressure tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the pressure tip wall resides approximately adjacent to the termination of the pressure sidewall; a suction tip wall that extends radially outwardly from the tip plate, traversing from the leading edge to the trailing edge such that the suction tip wall resides approximately adjacent to the termination of the suction sidewall; and one or more tip ribs that extend substantially between the pressure tip wall and the suction tip wall. 
         [0007]    These and other features of the present application will become apparent upon review of the following detailed description of the preferred embodiments when taken in conjunction with the drawings and the appended claims. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0008]    These and other objects and advantages of this invention will be more completely understood and appreciated by careful study of the following more detailed description of exemplary embodiments of the invention taken in conjunction with the accompanying drawings, in which: 
           [0009]      FIG. 1  is a partly sectional, isometric view of an exemplary gas turbine engine rotor blade mounted in a rotor disk within a surrounding shroud, with the blade having a tip in accordance with an exemplary embodiment of the present invention; and 
           [0010]      FIG. 2  is an isometric view of the blade tip as illustrated in  FIG. 1 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0011]    Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,  FIG. 1  depicts a portion of a turbine  10  of a gas turbine engine. The turbine  10  is mounted directly downstream from a combustor (not shown) for receiving hot combustion gases  12  therefrom. The turbine  10 , which is axisymmetrical about an axial centerline axis  14 , includes a rotor disk  16  and a plurality of circumferentially spaced apart turbine rotor blades  18  (one of which is shown) extending radially outwardly from the rotor disk  16  along a radial axis. An annular turbine shroud  20  is suitably joined to a stationary stator casing (not shown) and surrounds blades  18  for providing a relatively small clearance or gap therebetween for limiting leakage of combustion gases  12  therethrough during operation. 
         [0012]    Each blade  18  generally includes a dovetail  22  which may have any conventional form, such as an axial dovetail configured for being mounted in a corresponding dovetail slot in the perimeter of the rotor disk  16 . A hollow airfoil  24  is integrally joined to dovetail  22  and extends radially or longitudinally outwardly therefrom. The blade  18  also includes an integral platform  26  disposed at the junction of the airfoil  24  and the dovetail  22  for defining a portion of the radially inner flowpath for combustion gases  12 . It will be appreciated that the blade  18  may be formed in any conventional manner, and is typically a one-piece casting. 
         [0013]    It will be seen that the airfoil  24  preferably includes a generally concave pressure sidewall  28  and a circumferentially or laterally opposite, generally convex suction sidewall  30  extending axially between opposite leading and trailing edges  32  and  34 , respectively. The sidewalls  28  and  30  also extend in the radial direction between a radially inner root  36  at the platform  26  and a radially outer tip or blade tip  38 , which will be described in more detail in the discussion related to  FIG. 2 . Further, the pressure and suction sidewalls  28  and  30  are spaced apart in the circumferential direction over the entire radial span of airfoil  24  to define at least one internal flow chamber or channel for channeling cooling air through the airfoil  24  for the cooling thereof. Cooling air is typically bled from the compressor (not shown) in any conventional manner. 
         [0014]    The inside of the airfoil  24  may have any configuration including, for example, serpentine flow channels with various turbulators therein for enhancing cooling air effectiveness, with cooling air being discharged through various holes through airfoil  24  such as conventional film cooling holes  44  and trailing edge discharge holes  46 . 
         [0015]    As illustrated in  FIG. 2 , according to an exemplary embodiment of the present invention, blade tip  38  generally includes a tip plate  48  disposed atop the radially outer ends of the pressure and suction sidewalls  28  and  30 , where the tip plate  48  bounds internal cooling channel. The tip plate  48  may be integral to the rotor blade  18  or may be welded into place. A pressure tip wall  50  and a suction tip wall  52  may be formed on the tip plate  48 . Generally, the pressure tip wall  50  extends radially outwardly from the tip plate  48  (i.e., forming an angle of approximately 90° with the tip plate  48 ) and extends from the leading edge  32  to the trailing edge  34 . (Note that in some embodiments, the pressure tip wall  50  may form an angle with the tip plate  48  that is between 70° and 110°). The path of pressure tip wall  50  is adjacent to or near the termination of the pressure sidewall  28  (i.e., at or near the periphery of the tip plate  48  along the pressure sidewall  28 ). 
         [0016]    Similarly, the suction tip wall  52  extends radially outwardly from the tip plate  48  (i.e., forming an angle of approximately 90° with the tip plate  48 ) and extends from the leading edge  32  to the trailing edge  34 . (Note that in some embodiments, the suction tip wall  52  may form an angle with the tip plate  48  that is between 70° and 110°). The path of suction tip wall  52  is adjacent to or near the termination of the suction sidewall  30  (i.e., at or near the periphery of the tip plate  48  along the suction sidewall  30 ). 
         [0017]    Consistent with exemplary embodiments of the present invention, the height and width of the pressure tip wall  50  and/or the suction tip wall  52  may be varied depending on best performance and the size of the overall turbine assembly. As one of ordinary skill in the art will appreciate, the height and width of the pressure tip wall  50  and/or the suction tip wall  52  may be described in terms of their relative size in comparison to the radial length of the airfoil  24 . In preferred embodiments, the height of the pressure tip wall  50  and/or the suction tip wall  52  may be within the range of between about 0.1% to 10.0% of the radial height of the airfoil  24 . (Accordingly, put another way, if “HA” represents the approximate radial height of the airfoil and “HW” represents the approximate radial height of the pressure tip wall  50  or the suction tip wall  52 , then the ratio of HW/HA would be a value within the range of about 0.001 to 0.100.) More preferably, the height of the pressure tip wall  50  and/or the suction tip wall  52  may be within the range of between about 1% to 5% of the radial height of the airfoil  24 . Additionally, in preferred embodiments, the width of the pressure tip wall  50  and/or the suction tip wall  52  may be within the range of between about 0.1% to 5.0% of the radial height of the airfoil  24 . More preferably, the width of the pressure tip wall  50  and/or the suction tip wall  52  may be within the range of between about 0.5% to 2.5% of the radial height of the airfoil  24 . In addition, the pressure tip wall  50  and/or the suction tip wall  52  may extend in a continuous or intermittent manner, or may vary in height and width along its path, according to certain alternative embodiments. As shown, the pressure tip wall  50  and/or the suction tip wall  52  may be approximately rectangular in shape; other shapes are also possible. 
         [0018]    A tip mid-chord line  60  also is depicted on  FIG. 2 . As illustrated, the tip mid-chord line  60  is a reference line extending from the leading edge  32  to the trailing edge  34  that connects the approximate midpoints between the pressure tip wall  50  and the suction tip wall  52 . According to exemplary embodiments of the present application, one or more tip ribs  62  are formed on the blade tip  38 . As used herein, tip ribs  62  comprise narrow elongated protrusions that extend radially from the tip plate  48  (i.e., forming an angle of approximately 90° with the tip plate  48 ) and traverse across the tip plate  48  from the pressure tip wall  50  to the suction tip wall  52 . (Note that in some embodiments, the tip ribs  62  may form an angle with the tip plate  48  that is between 70° and 110°). In some embodiments, the present invention generally provides that the tip ribs  62  be configured such that a longitudinal axis  66  extending through each tip rib  62  forms an angle θ with the tip mid-chord line  60 , and that the angle θ fall within the following ranges. Preferably, angle θ is within a range of approximately 60°-120°, more preferably within a range of approximately 70°-110°, and optimally within a range of approximately 80°-100°. 
         [0019]    The number of tip ribs  62  may be vary depending upon best performance. In some embodiments, the tip ribs  62  will be approximately evenly spaced from the leading edge  32  to the trailing edge  34 . However, best performance may dictate that the spacing of the tip ribs  62  not be regular. The height and width of the tip ribs  62  may be varied depending on best performance and the size of the overall turbine assembly. In preferred embodiments, the height of the tip ribs  62  may be within the range of between about 0.1% to 10% of the radial height of the airfoil  24 . More preferably, the height of the tip ribs  62  may be within the range of between about 1.0% to 5% of the radial height of the airfoil  24 . In preferred embodiments, the width of the tip ribs  62  may be within the range of between about 0.1% to 5% of the radial height of the airfoil  24 . More preferably, the width of the tip ribs  62  may be within the range of between about 0.5% to 2.5% of the radial height of the airfoil  24 . The height and width of each tip rib  62  on a particular blade tip  38  may be approximately the same, though they may also vary depending on best performance. In addition, a particular tip rib  62  may be continuous or intermittent as it extends from the pressure tip wall  50  and the suction tip wall  52 . A particular tip rib  62  also may vary in height and width along its path, according to certain alternative embodiments and best performance. As shown, the tip ribs  62  may be approximately rectangular in shape; other shapes are also possible, such as a tip rib with rounded edges. In addition, in a preferred embodiment, the tip ribs  62  may extend radially past the height of either the pressure tip wall  50 , the suction tip wall  52 , or both. 
         [0020]    Further, as shown, the tip ribs  62  are straight. In some embodiments (not shown), the tip ribs  62  may be arcuate in shape. In such embodiments, the concave side of the tip rib  62  preferably will be on the upstream side of the rib. 
         [0021]    The present invention may be employed with any suitable manufacturing method. The pressure tip wall  50 , the suction tip wall  52 , and the tip ribs  62  may be formed, for example, by integral casting with the blade tip or complete blade, by electron-beam welding, by physical vapor deposition of material to a blade tip, or by brazing material. The present invention may be made with any suitable material, including the base metal or a dissimilar metallic or ceramic material, such as, for example, abradable TBC. 
         [0022]    In use, configurations of the pressure tip wall  50 , the suction tip wall  52 , and the one or more tip ribs  62 , according to the several embodiments discussed above, have been found to inhibit the flow of combustion gases through the gap between the turbine shroud  20  and the blade tip  38  by creating flow resistance therebetween. This, of course, increases the efficiency of the turbine engine because flow that leaks across the blade tip does not exert motive forces on the blade surfaces and accordingly is not providing work to the engine. In addition, it has been found that configurations according to the embodiments of the present invention could enhance the cooling characteristics that conventional systems (which typically include releasing cooling air through cooling holes located on the blade tip  38 ) provide to the blade tip region. Also, it has been found that configurations according to embodiments of the present invention generally enhance the aerodynamic performance of rotor blades. 
         [0023]    From the above description of preferred embodiments of the invention, those skilled in the art will perceive improvements, changes and modifications. Such improvements, changes and modifications within the skill of the art are intended to be covered by the appended claims. Further, it should be apparent that the foregoing relates only to the described embodiments of the present application and that numerous changes and modifications may be made herein without departing from the spirit and scope of the application as defined by the following claims and the equivalents thereof.