Patent Publication Number: US-9422816-B2

Title: Airfoil with hybrid drilled and cutback trailing edge

Description:
STATEMENT OF GOVERNMENT INTEREST 
     The present invention was developed, at least in part, with government funding pursuant to Contract No. N00019-02-C-3003 awarded by the United States Navy. The U.S. Government may have certain rights in this invention. 
    
    
     BACKGROUND 
     The present invention relates to fluid-cooled airfoils, and more particularly to fluid-cooled airfoils suitable for use with gas turbine engines. 
     Airfoils, such as those used in gas turbine engines, often operate in relatively hot environments. In order to help ensure air foil integrity, airfoils can utilize high temperature alloys, thermal barrier coatings, and cooling fluid delivery. However, known cooling schemes may be inadequate for some desired applications. Inadequate cooling fluid delivery can lead to spallation of coatings, and other wear or damage to the airfoil (e.g., crack formation), which may necessitate repair or replacement of the airfoil. Such a need for repair or replacement of an airfoil is costly and time-consuming. Therefore, it is desired to provide for improved fluid cooling for an airfoil, particularly at a trailing edge of the airfoil. 
     SUMMARY 
     An apparatus according to the present invention for use with a gas turbine engine includes an airfoil, a metering opening for metering a cooling fluid, a cutback slot configured to deliver the cooling fluid from the metering opening, and a cooling hole. The airfoil defines a trailing edge, opposite first and second faces, and a mean camber line. The cutback slot is defined along the first face of the airfoil adjacent to the trailing edge and is offset from the mean camber line of the airfoil. The cooling hole has an outlet that is located at the trailing edge and substantially aligned with the mean camber line of the airfoil. The cooling hole delivers a portion of the cooling fluid from the metering opening. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic cross-sectional view of a gas turbine engine. 
         FIG. 2  is a perspective view of an airfoil according to the present invention. 
         FIG. 3  is a cross-sectional view of a portion of the airfoil, taken along line  3 - 3  of  FIG. 2 . 
         FIG. 4  is an enlarged view of a portion of the airfoil, showing region IV of  FIG. 2 . 
         FIG. 5  is a schematic view of the airfoil, showing cooling flow and hot gas flow. 
         FIG. 6  is a flow chart of a method of making and using an airfoil according to the present invention. 
     
    
    
     DETAILED DESCRIPTION 
     In general, the present invention relates to a fluid-cooled airfoil having a film-cooling cutback slot located along a pressure face adjacent to the trailing edge and a convective-cooling hole extending to the trailing edge. A cooling fluid from a plenum is metered through a metering opening, and passes to the cutback slot to provide film cooling. A portion of the cooling fluid delivered to the cutback slot is directed through the cooling hole extending to the trailing edge to provide convective cooling to the airfoil. In that way, hybrid film cooling and convective cooling is provided at or near the trailing edge, which can help maintain regions of the trailing edge of the airfoil at or below suitable thermal operating limits. In one embodiment, an inlet of the hole extending to the trailing edge is located at or downstream from an upstream boundary of the cutback slot along the pressure face of the airfoil, and an outlet of the hole extending to the trailing edge is substantially aligned with a mean camber line of the airfoil. 
       FIG. 1  is a schematic cross-sectional view of a gas turbine engine  10  that includes a fan section  12 , a low-pressure compressor (LPC) section  14 , a high-pressure compressor (HPC) section  16 , a combustor section  18 , a high-pressure turbine (HPT) section  20 , and a low-pressure turbine (LPT) section  22 . A centerline C L  is defined by the engine  10 . A hot section  24  of the engine  10  is generally defined from the combustor section  18  afterward, including the HPT section  20  and the LPT section  22 . The illustrated embodiment of the gas turbine engine  10  is provided merely by way of example, and it should be recognized that the present invention applies to gas turbine engines of any configuration, such as low bypass ratio configurations. Those of ordinary skill in the art will understand the basic operation of gas turbine engines, and therefore further discussion here is unnecessary. 
       FIG. 2  is a perspective view of an airfoil  26  that defines a leading edge  28 , a trailing edge  30  downstream of the leading edge  28 , a pressure face  32 , and a suction face  34  (not visible in  FIG. 2 ; see  FIG. 3 ) located opposite the pressure face  32 . The airfoil  26  is suitable for use in the hot section  24  of the gas turbine engine  10 , and can be configured as either a blade or a stator. The airfoil  26  includes a plurality of cooling passages  36  at or near the trailing edge  30 . Additional cooling openings  38  of a known configuration can optionally be provided at upstream portions of the airfoil  26 . As shown in the illustrated embodiment, the cooling passages  36  are spaced apart ene from each other in a spanwise direction, and are located within a region defined between spanwise locations S 1  and S 2 . In one embodiment, the spanwise location S 1  is at approximately 30% of a span of the airfoil  26  and the spanwise location S 2  is at approximately 70-80% of the span of the airfoil  26 . The region defined between spanwise locations S 1  and S 2  can be selected to cover relatively high-temperature regions of the airfoil  26  near the trailing edge  30 . Moreover, limiting the region defined between spanwise locations S 1  and S 2  can help promote structural integrity of the airfoil  26  by omitting the cooling passages  36  at relatively high stress regions of the airfoil  26  (e.g., near a platform and tip). As shown in the illustrated embodiment, additional cutback slots  40  are located at the pressure face  32  adjacent to the trailing edge  30  at locations outside the region defined between spanwise locations S 1  and S 2 . It should be noted that the airfoil  26  can include a platform and a root, and in further embodiments can optionally include other features not specifically shown or described, such as a shroud. 
       FIG. 3  is a cross-sectional view of a portion of the airfoil  26 , taken along line  3 - 3  of  FIG. 2 . As shown in  FIG. 3 , a mean camber line  42  defines the mean thickness of the airfoil  26  between the pressure face  32  and the suction face  34 . In the illustrated embodiment, the trailing edge  30  is radiused, and has a diameter D 1 . A plenum  44  extends in at least partially in the spanwise direction inside the airfoil  26  for supplying a cooling fluid (e.g., bleed air). The plenum  44  can have a known configuration, and can act as a manifold to supply the cooling fluid to a number of cooling passages at various locations on the airfoil  26 . 
     Each of the cooling passages  36  (one is shown in  FIG. 3 ) includes a metering opening  46 , a cutback slot  48 , and a trailing edge cooling hole  50 . The metering opening  46  is fluidically connected to the plenum  44  to receive and meter cooling flows. The cutback slot  48  is located downstream from the metering opening  46 , and is configured to deliver cooling fluid to the pressure face  32  at an outlet defined between an upstream boundary  52  and a downstream boundary  54 . The downstream boundary  54  of the cutback slot  48  is located adjacent to and slightly upstream from the trailing edge  30 . The cutback slot  48  is generally offset from the mean camber line  42 . Additional details of the cutback slot  48  are discussed below. 
     The trailing edge cooling hole  50  extends from the cutback slot  48  to the trailing edge  30 , between an inlet  56  and an outlet  58 . In the illustrated embodiment, the inlet  56  of trailing edge cooling hole  50  is located essentially within the cutback slot  48 , that is, the inlet  56  is located approximately at or downstream of the upstream boundary  52  of the cutback slot  48  and at or upstream of the downstream boundary  54 . Furthermore, in the illustrated embodiment, the outlet  58  of the trailing edge cooling hole  50  is substantially aligned with the mean camber line  42  at the trailing edge  30 . The outlet  58  and other portions of the trailing edge cooling hole  50  has a substantially circular cross-section in the illustrated embodiment. In alternative embodiments, other shapes of the outlet  58  are possible, such as an elliptical or “racetrack” shape with a major axis arranged in the spanwise direction. The outlet  58  has a diameter (or width) D 2 . In one embodiment, the diameter D 1  of the trailing edge  30  is at least approximately three times larger than the diameter D 2  of the outlet  58 . Having the diameter D 1  significantly larger than the diameter D 2  helps promote structural integrity of the trailing edge  30 . 
     Although in the illustrated embodiment only a single trailing edge cooling hole  50  extends from each cutback slot  48 , in further embodiments multiple trailing edge cooling holes  50  can extend from a given cutback slot  48 . For example, multiple trailing edge cooling holes  50  can extend from a given cutback slot  48  at different angles relative to the centerline C L  and each have separate inlets  56 . Alternatively, multiple trailing edge cooling holes  50  extending from a given cutback slot  48  could share a common inlet  56 . 
       FIG. 4  is an enlarged view of a portion of the airfoil  26 , showing region IV of  FIG. 2 . In the embodiment illustrated in  FIG. 4 , the cutback slots  48  each have a diverging shape at their respective outlets. The cutback slots  48  each define an outlet area that is substantially larger than that of either the inlet  56  or the outlet  58  of the corresponding trailing edge cooling hole  50 . Furthermore, as shown in  FIG. 4 , the cooling passages  36  extend substantially axially with respect to the centerline C L  of the engine  10 , and the trailing edge cooling holes  50  are each substantially aligned with a corresponding one of the cutback slots  48  at a given spanwise location. In alternative embodiments, the cooling passages  36  can have different orientations as desired for particular applications. For example, in an alternative embodiment the cutback slot  48  and the trailing edge cooling hole  50  of any of the cooling passages  36  can extend at different angles with respect to the centerline C L . 
       FIG. 5  is a schematic view of the airfoil  26 , showing a cooling fluid flow  60  and hot gas flows  62 . During operation, the hot gas flows  62  pass along the pressure face  32  and the suction face  34  of the airfoil  26 , and continue past the trailing edge  30 . The relatively cool cooling fluid flow  60  is supplied by the plenum  44  to the cooling passages  36  (only one cooling passage  36  is shown in  FIG. 5 ). The cooling fluid flow  60  is delivered to the cutback slot  48 , and a first portion  60 A of the cooling fluid flow  60  is exhausted from the cutback slot  48  at the pressure face  32  of the airfoil  26  to provide film cooling at or near the trailing edge  30 . Film cooling tends to create a layer of relatively cool fluid between the hot gas flows  62  and surfaces of the airfoil  26  in order to help keep the airfoil  26  cool. A second portion  60 B of the cooling fluid flow  60  is delivered by the trailing edge cooling hole  50 , and the second portion  60 B is diverted from the cutback slot  48  and exhausted from the trailing edge  30  to provide convective cooling at or near the trailing edge  30 . Convective cooling allows thermal energy from the airfoil  26  to be absorbed by the cooling fluid flow  60  and thereby removed and exhausted to the hot gas flows  62 . 
     The second portion  60 B of the cooling fluid flow  60  also provides aerodynamic benefits by helping to straighten fluid flows at or near the trailing edge  30  of the airfoil  26 . Moreover, by exhausting the second portion  60 B of the cooling fluid flow  60  at the trailing edge  30  along the mean camber line  42 , the relatively high mixing losses typically associated with pressure face and suction face cooling flows are avoided. 
       FIG. 6  is a flow chart of a method of making and using the airfoil  26 . First, the airfoil  26  is created using a casting process (step  100 ). During casting, one or more of cutback slots  48  are defined, which can be accomplished using casting cores in a known manner. After the cutback slots  48  are defined, one or more trailing edge cooling holes  50  are drilled in the airfoil  26  (step  102 ). Drilling can be performed using electric discharge machining (EDM), laser drilling, or other suitable processes. When multiple trailing edge cooling holes  50  are desired, they can be drilled simultaneously or sequentially. A thermal barrier coating (TBC) is also applied to the airfoil  26  (step  104 ). In one embodiment, the TBC is applied subsequent to drilling of the trailing edge cooling holes. However, with some drilling methods, such as laser drilling, the TBC could alternatively be applied prior to drilling. Furthermore, in some embodiments, use of the TBC can be omitted entirely. Lastly, for each cooling passage  36 , a cooling fluid is supplied and metered when the airfoil  26  is in operation (step  106 ), and the metered cooling fluid is delivered to the cutback slot  48  to provide film cooling (step  108 ). A portion of the cooling fluid delivered to the cutback slot  48  is diverted for delivery by the trailing edge cooling holes  50  (step  110 ). 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. For example, the present invention can be utilized in conjunction with any number of additional cooling features, such as additional cooling passages of a known configuration. Moreover, trailing edge cooling holes can be drilled into existing airfoils with cutback slots as part of a repair or retrofit operation according to the present invention.