Abstract:
A gas turbine engine component includes a structure having a surface configured to be exposed to a hot working fluid. The surface includes a recessed pocket that is circumscribed by an overhang. At least one cooling groove is provided by the overhang.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/902,497, which was filed on Nov. 11, 2013 and is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    This disclosure relates to a gas turbine engine. More particularly, the disclosure relates to a tip cooling configuration for an airfoil. 
         [0003]    Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
         [0004]    Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. 
         [0005]    In the pursuit of ever higher efficiencies, gas turbine manufacturers have long relied on higher and higher turbine inlet temperatures to provide boosts to overall engine performance. In typical modern engine applications the gas path temperatures within the turbine exceed the melting point of the component constituent materials. Due to this, dedicated cooling air is extracted from the compressor and used to cool the gas path components in the turbine incurring significant cycle penalties. 
         [0006]    Turbine blades typically include internal cooling passages. Film cooling holes communicate cooling fluid from the cooling passages to high temperature areas on the exterior surface of the turbine blade that may experience undesirably high temperatures. One high temperature area is the tip of the airfoil. 
       SUMMARY 
       [0007]    In one exemplary embodiment, a gas turbine engine component includes a structure having a surface configured to be exposed to a hot working fluid. The surface includes a recessed pocket that is circumscribed by an overhang. At least one cooling groove is provided by the overhang. 
         [0008]    In a further embodiment of the above, the cooling channel exits through a continuous channel into the recessed pocket. 
         [0009]    In a further embodiment of any of the above, the cooling channel exits through discontinuous channels into the recessed pocket. 
         [0010]    In a further embodiment of any of the above, the component includes at least one discrete hole that is in fluid communication with the groove and is configured to provide a cooling fluid to the pocket. 
         [0011]    In a further embodiment of any of the above, the structure is an instrumentation probe. 
         [0012]    In a further embodiment of any of the above, the structure is an airfoil. 
         [0013]    In a further embodiment of any of the above, the airfoil includes a cast first portion, and a second portion is secured to the first portion, the second portion providing the overhang. 
         [0014]    In a further embodiment of any of the above, the second portion is additively manufactured. 
         [0015]    In a further embodiment of any of the above, the overhang circumscribes the pocket. 
         [0016]    In a further embodiment of any of the above, the overhang includes a lip that provides an interior perimeter of the pocket. The groove is provided between the overhang and the end wall. The groove is bounded by the lip. 
         [0017]    In a further embodiment of any of the above, the overhang substantially encloses the groove and provides an exit that fluidly interconnects the groove with the pocket. 
         [0018]    In a further embodiment of any of the above, the exit is provided radially between the lip and the end wall. 
         [0019]    In a further embodiment of any of the above, the pocket is teardrop-shaped. 
         [0020]    In a further embodiment of any of the above, the overhang and an adjacent wall encloses the groove. 
         [0021]    In another exemplary embodiment, a method of manufacturing a turbine blade airfoil, includes the step of forming a structure having a surface configured to be exposed to a hot working fluid, forming a surface comprising a recessed pocket, forming an overhang that circumscribes the recessed pocket which includes at least one cooling groove provided by the overhang, and using an additive manufacturing process to create a negative for casting of features for at least one of the steps. 
         [0022]    In a further embodiment of the above, wherein the forming steps are performed by directly successively adding layers of metal powder joined by local directed energy such as direct laser metal sintering, selective laser metal melting, or electron beam melting. The using step is replaced by an injection molded ceramic core or stamped refractory metal negative for casting of features for at least one of the forming steps. The using step further includes successively adding layers of metal powder to a partially cast component for construction of at least one of the forming steps. 
         [0023]    In a further embodiment of the above, the method includes additively manufacturing at least one core that provides a cavity that has an airfoil shape that corresponds to the airfoil. The forming step includes casting the airfoil within the cavity. 
         [0024]    In a further embodiment of any of the above, the forming step includes casting a first airfoil portion, and additively manufacturing a second airfoil portion onto the first airfoil portion. The second airfoil portion provides the overhang. 
         [0025]    In another exemplary embodiment, a method of manufacturing a gas turbine engine component, includes the steps of forming step a first airfoil portion, and additively manufacturing a second airfoil portion onto the first airfoil portion, the second airfoil portion including a recessed pocket that is circumscribed by an overhang, and at least one cooling groove provided by the overhang. 
         [0026]    In a further embodiment of any of the above, the first air foil portion is cast. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0028]      FIG. 1A  is a perspective view of the airfoil having the disclosed cooling passage. 
           [0029]      FIG. 1B  is a plan view of the airfoil illustrating directional references. 
           [0030]      FIG. 2A  is an enlarged perspective view of an airfoil tip. 
           [0031]      FIG. 2B  is a cross-sectional view through the tip taken along lines  2 B- 2 B in  FIG. 2A . 
           [0032]      FIG. 3A  is a top elevational view of the airfoil tip shown in  FIG. 2A . 
           [0033]      FIG. 3B  is a cross-sectional perspective view of the tip taken along line  3 B- 3 B of  FIG. 3A . 
           [0034]      FIG. 3C  is a cross-sectional end view of the tip shown in  FIG. 3B . 
           [0035]      FIG. 4  is a schematic view of a mold used in forming the airfoil shown in  FIG. 3B . 
           [0036]      FIG. 5  is a schematic view of a gas turbine engine component having an overhang arranged above a groove. 
           [0037]      FIG. 6  is a schematic view of a gas turbine engine component having an overhang enclosing a groove. 
       
    
    
       [0038]    The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
       DETAILED DESCRIPTION 
       [0039]    The disclosed cooling configuration may be used in various gas turbine engine applications. A gas turbine engine uses a compressor section that compresses air. The compressed air is provided to a combustor where the compressed air and fuel is mixed and burned. The hot combustion gases pass over a turbine to provide work that may be used for thrust or driving another system component. Many of the engine components, such as blades, vanes, combustor and exhaust liners, blade outer air seals and instrument probes, are subjected to very high temperatures such that cooling may become necessary. The disclosed cooling configuration and manufacturing method may be used for any gas turbine engine component. For exemplary purposes, a turbine blade  10  is described. 
         [0040]    Referring to  FIGS. 1A and 1B , a root  12  of each turbine blade  10  is mounted to the rotor disk  16 . The turbine blade  10  includes a platform  14 , which provides the inner flow path, supported by the root  12 . An airfoil  18  extends in a radial direction R from the platform  14  to a tip  28 . It should be understood that the turbine blades may be integrally formed with the rotor such that the roots are eliminated. In such a configuration, the platform is provided by the outer diameter of the rotor. The airfoil  18  provides leading and trailing edges  20 ,  22 . The tip  28  is arranged adjacent to a blade outer air seal  30 . 
         [0041]    The airfoil  18  of  FIG. 1B  somewhat schematically illustrates exterior airfoil surface extending in a chord-wise direction C from a leading edge  20  to a trailing edge  22 . The airfoil  18  is provided between pressure (typically concave) and suction (typically convex) wall  24 ,  26  in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple turbine blades  10  are arranged circumferentially in a circumferential direction A. The airfoil  18  extends from the platform  14  in the radial direction R, or spanwise, to the tip  28 . 
         [0042]    The airfoil  18  includes a cooling passage  32  provided between the pressure and suction walls  20 ,  22 . The exterior airfoil surface  34  may include multiple film cooling holes (not shown) in fluid communication with the cooling passage  32 . 
         [0043]    Referring to  FIGS. 2A-3C , the tip  28  includes an end face  36 , which is configured to be adjacent to the BOAS. A pocket  38  is provided in the tip  28 . An overhang  42  circumscribes the pocket  38  and provides the end face  36 . The overhang  42  forms an interior perimeter  44  arranged within the pocket  38 . As best shown in  FIGS. 3B-3C , the overhang  42  includes a radially inwardly extending lip  46 , which forms a groove  50  that is substantially enclosed by the overhang  42 . An exit  48  is provided between the lip  46  and the end wall  39 . 
         [0044]    At least one cooling hole  40 , round or shaped, extend through the end wall  39  in a generally radial direction to fluidly interconnect the cooling passage  32  and the groove  50 . The holes  40  can be oriented in other directions, if desired. An impingement cooling flow is provided through the at least one hole  40  into the groove  50  and onto the overhang  42 , which cools the end face  36 . Cooling fluid within the groove is permitted to pass through the exit  48  and into the pocket  38 . 
         [0045]    The at least one discrete holes lie around the tip cap and are angled to the most optimal impingement location along the tip region. The holes would be angled such that they impinge on the interior of the cavity while balancing degradation effects of their impingement angle. The post impingement air pressurizes the cavity. The air then ejects through the blade tip such that the pocket  38  acts as a traditional blade tip film cooling. 
         [0046]    The cooling configuration employs relatively complex geometry that cannot be formed by traditional casting methods. To this end, additive manufacturing techniques may be used in a variety of ways to manufacture an airfoil with the disclosed cooling configuration. In one example, as schematically illustrated in  FIG. 3C , a first portion  60  of the airfoil  18  may be formed by a typical casting technique. The tip  28  of the airfoil may be formed by additively manufacturing a second portion  62  in which the remaining tip is deposited directly on to the casting portion  60 . For example, the cast blade is placed into a fixture within a powder-bed additive machine (such as an EOS  280 ) and the last  10 % of the blade is directly additively manufactured. 
         [0047]    Other manufacturing techniques are schematically illustrated in  FIG. 4 . This core could be constructed using a variety of processes such as photo-polymerized ceramic, electron beam melted powder refractory metal, or injected ceramic based on an additively built disposable core die. The core and/or shell molds for the airfoils are first produced using a layer-based additive process such as LAMP from Renaissance Systems. Further, the core could be made alone by utilizing EBM of molybdenum powder in a powder-bed manufacturing system. 
         [0048]    A ceramic outer mold  52  and interior core mold  54  may be additively manufactured separately or as one piece to form a cavity  58  providing an airfoil shape. Molten metal is cast into cavity  58  to form the airfoil  18 . Pins  56  interconnect the outer mold  52  and interior core mold  54  to provide the correspondingly shaped cooling holes. 
         [0049]      FIG. 5  illustrates a component  110  having an overhang  142  spaced from a wall  139  to provide a groove  150 . A cooling hole  140  is in fluid communication with the groove  150 , which provides an exit  148  without a lip. 
         [0050]    Another component  210  is shown in  FIG. 6 . The component  210  includes a groove  250  enclosed by the overhang  242  and the wall  239 . A cooling hole  240  communicates cooling fluid to the groove  250 . Fluid enters the pocket  238  through exits  248 , which may be provided by slots, for example. 
         [0051]    The cooling configuration provides increased engine efficiency through a realizable turbine blade cooling configuration with increased effectiveness of blade tip cooling. The shaped channel design provides cold wall surface area allowing for internal convection, increasing effectiveness over a normal tip cooling configuration. 
         [0052]    It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention. 
         [0053]    Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0054]    Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.