• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a turbine blade (100) for a gas turbine. The turbine blade (100) includes a platform (108), an airfoil (104) extending radially from the platform (108) and a plurality of in the airfoil (104) and near an exterior surface (122) of the airfoil (104). defined cooling channels (120). Each of the cooling channels (120) may include a radially inner portion (132) having a first cross-sectional area and at least one radially outer portion (134) having a second cross-sectional area, the first cross-sectional area being greater than the second cross-sectional area. The cooling channels (120) are formed as a groove in the cover layer covered by a base portion of the airfoil (104).
    公开号:CH707844B1
    申请号:CH00353/14
    申请日:2014-03-10
    公开日:2018-02-28
    发明作者:Paul Lacy Benjamin
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    description
    TECHNICAL FIELD OF THE INVENTION The present application and the pending patent relate primarily to gas turbine engines, and more particularly, to gas turbine engine blade cooling ducts for providing improved cooling at high operating temperatures.
    Background of the Invention In a gas turbine, hot combustion gases flow from one or more burners through a transition piece and along a hot gas path. A number of turbine stages are typically arranged sequentially along the hot gas path such that the combustion gases flow through first stage vanes and blades, and then through later stage vanes and blades of the turbine. In this way, the turbine blades are exposed to high temperatures resulting from the combustion gases flowing along the hot gas path. Since the efficiency of a gas turbine is dependent on its operating temperatures, there is a constant need for components along the hot gas path, such as those shown in FIGS. Turbine blades, increasingly higher temperatures without failure or shortening of the operating life can withstand.
    Certain turbine blades, particularly those of later turbine stages, may include a number of cooling holes that extend radially through the turbine blades. In this way, the cooling holes may be a cooling fluid, such as a cooling fluid. Air, through the turbine blade to exchange heat to keep the temperature of the turbine blade within a permissible range. According to a known cooling hole design, the turbine blade may include a number of long straight cooling holes produced by electrolytic machining with a drill pipe otherwise known as "STEM drilling". Although such an arrangement may provide adequate cooling for the turbine blade in certain applications, cooling holes made by conventional STEM drilling are limited to a straight path through the turbine blade. As a result, the three-dimensional shape of the turbine blade is also limited due to the need to accommodate the straight cooling holes extending therethrough. Further, the straight cooling holes produced by STEM drilling have a constant diameter, and thus do not satisfy the variation in the cooling requirements along the radial length of the turbine blade. In particular, as a result of the constant diameter, an undesirable amount of heat may be transferred to the cooling fluid before it reaches a tip region of the turbine blade where the cooling requirement is greater.
    Thus, there is a desire for an improved turbine blade having a cooling configuration that resists high temperatures along the hot gas path of a gas turbine engine. In particular, such a cooling configuration may allow the turbine blade to have various complex three-dimensional shapes or a twist for improved aerodynamics. Such a cooling configuration may also accommodate the variations in cooling requirements along the radial length of the turbine blade for efficient cooling. Finally, such a cooling configuration may reduce the amount of airflow required to cool the turbine blade while increasing the overall efficiency of the gas turbine.
    Brief Description of Embodiments of the Invention The present application and the resulting patent thus provide a turbine blade for a gas turbine. The turbine bucket includes a platform, an airfoil extending radially from the platform, and a number of cooling channels defined in the airfoil and in the vicinity of an outer surface of the airfoil. Each of the cooling passages includes a radially inner portion having a first cross-sectional area and at least one radially outer portion having a second cross-sectional area defined by branched cooling passages, the first cross-sectional area being greater than the second cross-sectional area.
    At least one of the cooling passages of the above-mentioned turbine blade may have more than one radially outer portion.
    At least one of the cooling passages of each turbine blade mentioned above may have a radially intermediate portion extending between the radially inner portion and the at least one radially outer portion, wherein the at least one radially intermediate portion may have a third cross-sectional area and the third cross-sectional area is greater than the second cross-sectional area and smaller than the first cross-sectional area.
    The airfoil of each turbine blade mentioned above includes a base portion and a cover layer extending above the base portion, each of the cooling channels being defined by a groove formed in the base portion and over which the cover layer extends.
    The cover layer of each turbine blade mentioned above may include a first thickness above the radially inner portion and a second thickness above the radially outer portion, wherein the first thickness is greater than the second thickness.
    At least one of the cooling channels of each turbine blade mentioned above may extend radially along the outer surface of the airfoil from the platform to one or more outlets defined in a tip shroud of the turbine blade.
    At least one of the cooling channels of each turbine blade mentioned above may extend radially along the outer surface of the airfoil from the platform to one or more outlets defined in the outer surface of the airfoil and positioned between the platform and a tip shroud of the turbine blade ,
    The turbine blade of each type mentioned above may further comprise a stem extending radially from the platform and away from the airfoil, and at least one feed channel defined in the stem, the feed channel communicating with one or more of the cooling channels at one Junction can be connected.
    The feed channel of each turbine blade mentioned above may be in communication with one or more of the radially inner portions at the joint, and the joint may be positioned in the platform.
    The feed channel of each turbine blade mentioned above may communicate with one or more of the radially inner portions at the joint, and the joint may be positioned in the airfoil between the platform and a tip shroud of the turbine blade.
    The feed channel of each turbine blade mentioned above may have a fourth cross-sectional area, and the fourth cross-sectional area may be larger than the first cross-sectional area.
    These and other features and improvements of the present application will become apparent to those skilled in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims.
    Short description of the drawings [0017]
    Fig. 1 is a schematic representation of a gas turbine with a compressor, a burner and a turbine.
    Fig. 2 is a schematic representation of a portion of a turbine, as it can be used ver in the gas turbine of Fig. 1, which represents a number of turbine stages.
    Fig. 3 is a front elevational view of a prior art turbine blade as may be used in the turbine of Fig. 2, illustrating a number of cooling holes illustrated by hidden lines.
    FIG. 4 is a plan view of the turbine blade of FIG. 3. FIG.
    FIG. 5 is a front elevational view of one embodiment of a turbine blade as may be described herein illustrating a number of cooling holes illustrated by hidden lines. FIG.
    Fig. 6 is a front elevational view of another embodiment of a turbine blade as may be described herein illustrating a number of cooling holes illustrated by hidden lines.
    7 is a cross-sectional top view of a portion of a turbine blade as may be described herein illustrating a cooling channel defined in the turbine blade.
    8 is a cross-sectional top view of a portion of a turbine blade as may be described herein illustrating a cooling channel defined in the turbine blade.
    Detailed Description of the Invention In the drawings, wherein like reference numerals designate the same elements throughout the several views, FIG. 1 illustrates a schematic view of a gas turbine engine 10 as may be used herein. The gas turbine 10 may include a compressor 15. The compressor 15 compresses an incoming air stream 20. The compressor 15 delivers the compressed air stream 20 to a burner 25. The burner 25 mixes the compressed air stream 20 with a pressurized fuel stream 30 and burns the mixture to produce a stream of combustion gases 35 , Although only a single burner 25 is shown, the gas turbine 10 may include any number of burners 25. The flow of combustion gases 35 is then delivered to a turbine 40. The flow of combustion gases 35 drives the turbine 40 to produce mechanical work. The mechanical work generated in the turbine 40 drives the compressor 15 via a shaft 45 and an external load 50, e.g. an electric generator and the like. Other embodiments and other components can be used herein.
    Gas turbine 10 may utilize natural gas, various types of synthesis gas, and / or other types of fuels. Gas turbine 10 may be one of a number of different gas turbines offered by General Electric Company of Schenectady, New York, including, but not limited to, those of a high performance Series 7 and 9 gas turbine, and the like. The gas turbine 10 may have different arrangements, and may use other types of components. Other types of gas turbines may also be used herein. Several gas turbines, other types of turbines, and other types of power generation technology may also be used herein. Although the gas turbine engine 10 is illustrated herein, the present application is applicable to any type of turbomachinery.
    FIG. 2 illustrates a schematic view of a portion of the turbine 40 having a number of stages 52 positioned in a hot gas path 54 of the gas turbine engine 10. A first stage 56 may include a number of circumferentially spaced vanes 58 and first stage buckets 60. Likewise, a second stage 62 may include a plurality of circumferentially spaced vanes 64 and second stage buckets 66. Further, a third stage 68 may include a number of circumferentially spaced vanes 70 and third stage buckets 72. Although the portion of the turbine 40 is shown as having three stages 52, the turbine 40 may include any number of stages 52.
    3 illustrates a front elevational view of one of the second stage rotor blades 66 of the turbine 40. As is known, the blade 66 includes an airfoil 76, a stem 78, and a platform 80 disposed between the airfoil 76 and the stem 78 are arranged. The airfoil 76 extends radially from the platform 80 to a tip 82 positioned at a tip end 84 of the blade 66. The tip shroud 82 may be integrally formed with the airfoil 76. The stem 78 may extend radially from the platform 80 to a root end 86 of the blade 66 such that the platform 80 substantially defines a juncture between the airfoil 76 and the stem 78. As illustrated, the platform 80 may be configured to be substantially planar and substantially horizontal when the bucket 66 is positioned in the turbine 40 for use. The shaft 78 may be used to define a foot structure, such as a foot structure. a dovetail configured to secure the blades 66 to a turbine disk of the turbine 40. During operation of the gas turbine 10, the flow of combustion gases 35 travels along the hot gas path 54 and over the platform 80, which forms the radially inner boundary of the hot gas path 54. As a result, the flow of combustion gases 35 is directed against the airfoil 76 of the blade 66 and thus the surfaces of the airfoil 76 are exposed to very high temperatures.
    As shown in FIGS. 3 and 4, the bucket 66 includes a number of cooling holes 88 (shown as hidden lines) defined in the airfoil 76. Each cooling hole 88 extends radially out of the platform 80 to an outlet 90 defined in the tip shroud 82 at the tip end 84 of the blade 66. The cooling holes 88 may be made by conventional STEM drilling and thus may have a substantially circular cross-sectional shape and a constant diameter along the length of the cooling holes 88 have. The blade 66 may also include a number of feed holes 92 (shown in phantom lines) defined in the stem 78. Each feed hole 92 may extend radially from an inlet 94 defined in the stem 78 at the root end 86 of the blade 66 to the platform 80. As shown, each feed hole 92 may be in direct communication with one of the cooling holes 88 at a joint 96 positioned in the platform 80. The feed holes 92 may also be made by conventional STEM drilling and thus have substantially a circular cross-section and a constant diameter along the length of the feed holes 92. During operation of the gas turbine 10, a cooling fluid such as e.g. Bleed air are passed from the compressor 15 into the feed holes 92 by means of the inlets 94 and can then pass through the cooling holes 88 and exit the blade 66 via the outlets 90. As a result, heat from the blade 66, and particularly the airfoil 76, may transfer to the cooling fluid as it passes through the cooling holes 88 and is then directed into the hot gas path 54 at the tip end 84 of the blade 66.
    FIG. 5 illustrates a front elevational view of one embodiment of a turbine blade 100 as described herein. FIG. The blade 100 may be in a later stage, such as the second stage 62 of the turbine 40, the gas turbine 10 are used. In a similar manner, the bucket 100 may be used in the third stage 68 or any other stage of the turbine 40. The bucket 100 includes an airfoil 104, a shaft 106, and a platform 108 disposed between the airfoil 104 and the shaft 106. The airfoil 104 extends radially from the platform 108 to a tip shroud 112 positioned at a tip end 104 of the bucket 104. The tip shroud 112 may be integrally formed with the airfoil 104. The shaft 106 may extend radially downwardly from the platform 108 to a root end 118 of the blade 100 such that the platform 108 substantially defines a juncture between the airfoil 104 and the shaft 106. As illustrated, the platform 108 may be configured to be substantially planar and substantially horizontal when the bucket 100 is positioned in the turbine 40 for use. The shaft 106 may be shaped to have a foot structure, such as a foot structure. defines a dovetail which is used to attach the blade 100 to a
    Turbine disk of the turbine 40 is designed. Other arrangements of the airfoil 104, shaft 106, and platform 108 may be used.
    As shown in FIG. 5, the bucket 100 includes a number of cooling channels 120 (shown by hidden lines) defined in the airfoil 104 proximate an outer surface 122 of the airfoil 104. The cooling channels 120 may extend radially along the outer surface 122 of the airfoil 104 from the platform 108 to the tip end 114 of the blade 100. In particular, one or more cooling channels 120 may extend radially along the outer surface 122 of the airfoil 104 from the platform 108 to one or more outlets 126 defined in the tip shroud 112 at the tip end 114 of the bucket 100. In addition, one or more cooling channels 120 may extend radially along the outer surface 122 of the airfoil 104 and outlets 128 radially defined between the platform 108 and the tip shroud 112. Likewise, one or more of the cooling channels 120 each extend radially along the outer surface 122 of the airfoil 104 from the platform 108 to one or more outlets defined along the blade 100. For example, in certain aspects, one or more of the cooling channels 120 may extend radially to one or more outlets defined by one or more through holes along the airfoil 104. In some aspects, one or more of the cooling channels 120 may extend radially along the outer surface 122 of the airfoil 104 and follow the contour of the outer surface 122 of the airfoil 104. In other aspects, one or more of the cooling channels 120 may extend axially along the outer surface 122 of the airfoil 104 to one or more outlets defined in the airfoil 104. For example, in certain aspects, one or more of the cooling channels 120 may extend axially to one or more outlets defined by one or more through holes positioned along the airfoil. In still other aspects, one or more of the cooling channels 120 may each axially reverse to form a serpentine pattern or any other pattern for optimizing cooling of the blade 100. In some aspects, one or more of the cooling channels 120 may extend to one or more outlets defined by one or more through holes positioned along the trailing edge of the airfoil 104.
    One or more of the cooling channels 120 may each have a cross-sectional area that varies along the radial length of the cooling channel 120. The cross-sectional area of the cooling passage 120 may gradually vary in a tapered manner along the radial length of the cooling passage 120. Alternatively, the cross-sectional area of the cooling passage 120 may suddenly vary in a stepwise manner along the radial length of the cooling passage, as shown in FIG. In particular, one or more of the cooling channels 120 may have a cross-sectional area in the vicinity of the platform 108 that is larger than a cross-sectional area of the cooling channel 120 near the tip end 114 of the blade 100.
    In some aspects, one or more of the cooling channels 120 may each have a radially inner portion 132 having a cross-sectional area that is constant along the radial length of the radially inner portion 132 and a radially outer portion 134 having a second cross-sectional area along the radial length of the radially outer portion 134 is constant. According to the invention, the first cross-sectional area is greater than the second cross-sectional area. The radially inner portion 132 communicates with the radially outer portion 134 at a junction 136 positioned between the platform 108 and the tip end 114 of the bucket 100. Alternatively, one or more of the cooling channels 120 may include a radially intermediate portion 138 positioned between the radially inner portion 132 and the radially outer portion 134. In certain aspects, the radially intermediate portion 138 may have a third cross-sectional area that is constant along the radial length of the radially intermediate portion 138. In such aspects, the third cross-sectional area may be smaller than the first cross-sectional area and larger than the second cross-sectional area. The radially intermediate portion 138 may be in direct communication with the radially inner portion 132 at a junction 142 and in direct communication with the radially outer portion 134 at a junction 144. As illustrated, each of the junctions 136, 142, 144 between the various sections 132, 134, 138 may have a step transition from one cross-sectional area to another cross-sectional area. Alternatively, each of the junctions 136, 142, 144 between the various sections 132, 134, 138 may have a tapered transition from one cross-sectional area to another cross-sectional area. Other arrangements may be used in which one or more cooling channels may each include additional spaced intermediate portions positioned between the radially inner portion 132 and the radially outer portion 134, each additional spaced intermediate portion being a particular and different one Has cross-sectional area.
    According to the illustration in FIG. 5, according to the invention, one or more of the cooling channels 120 has at least one radially outer section 134 with a plurality of branches standing in direct connection with the radially inner section 132 at the connection point 136. In this way, the cooling channel 120 defines a tree-like shape, whereby the radially outer portions 134 of the radially inner portion 132 at the connection point 136 abzu¬ gene. Further according to the invention, the first cross-sectional area of the radially inner portion 132 is larger than the second cross-sectional area of each of the radially outer portions 134. Similarly, one or more of the cooling channels 120 may each have a number of intermediate portions 138 in communication with the radially inner portion 132 at the junction 142, which forms a branched tree-like shape. In some aspects, the first cross-sectional area of the radially inner portion 132 may be larger than the third cross-sectional area of each of the intermediate portions 138. Further, one or more of the cooling channels 120 may each have a number of radially outer portions 134 in direct communication with one of the intermediate portion 134 at the junction 144, which also forms a tree-like shape. In certain aspects, the third cross-sectional area of the intermediate portion 138 may be greater than the second cross-sectional area of each of the radially outer portions 134. Other arrangements may be used in which one or more cooling channels may each include additional spaced intermediate portions positioned between the radially inner portion 132 and the radially outer portion 134, each additional spaced intermediate portion being a particular and different one Has cross-sectional area.
    The bucket 100 may also include one or more feed lines 148 (shown in hidden lines) defined in the bucket 100. As shown in FIG. 5, the one or more feed channels 148 may extend radially from an inlet 152 defined in the stem 106 at the foot end 118 of the bucket 100 to the platform 108. The one or more feed channels 148 may each be in direct communication with one or more of the cooling channels 120 at the joint 154 defined in the platform 108. In particular, the one or more feed channels 148 may each be in direct communication with one or more of the radially inner portions 132 at the joint 154. In some aspects, the one or more feed channels 148 may each have a fourth cross-sectional area that is constant along the radial length of the feed channel 148. In such aspects, the fourth cross-sectional area may be greater than the first cross-sectional area of each of the one or more radially inner portions 132.
    6 is a front elevational view of another embodiment of a turbine bucket 200 described herein. The bucket 200 includes various elements that correspond to the elements described above with reference to the bucket 100, which elements are designated by like reference numerals in FIG are not described in more detail herein. The blade 200 may be used in a later stage, such as in the US. the second stage 62 of the turbine 40 of the gas turbine 10 are used. Similarly, the bucket 200 may be used in the third stage 68 or any other stage of the turbine 40.
    As shown in FIG. 6, the bucket 200 includes a number of cooling channels 220 (shown by hidden lines) defined in the airfoil 204 near the outer surface 222 of the airfoil 204. The cooling channels 220 may extend radially along the outer surface 222 of the airfoil 204 from a position near the platform 208 to the tip end 214 of the bucket 200. In particular, one or more of the cooling channels 220 may each extend radially along the outer surface 222 of the airfoil 204 from the position near the platform 208 to one or more outlets 226 defined in a tip shroud 212 at the tip end 214 of the blade 200 , In addition, one or more of the cooling channels 220 may extend radially along the outer surface 222 of the airfoil 204 from the position near the platform 208 to one or more outlets 228 defined in the outer surface 222 of the airfoil 204 and radially between the platform 208 and the tip shroud 212 are positioned. In some aspects, one or more of the cooling channels 220 may include one or more radially inner portions 132 and one or more radially outer portions 134 that correspond to the elements described above with reference to the blade 100. Likewise, one or more of the cooling channels 220 may each extend radially along the outer surface 222 of the airfoil 204 from the platform 208 to one or more outlets defined anywhere along the airfoil 200. In some aspects, one or more of the cooling channels 220 may extend radially along the outer surface 222 of the airfoil 204 and follow the contour of the outer surface 222 of the airfoil 204. In other aspects, one or more of the cooling channels 220 may each extend axially along the outer surface 222 of the airfoil 204. For example, in certain aspects, one or more of the cooling channels 220 may extend axially to one or more outlets defined by one or more through holes positioned along the airfoil 204. In still other aspects, one or more of the cooling channels 220 may reverse in the axial direction or form serpentine patterns or any other pattern for optimizing the cooling of the blade 200. In such aspects, one or more of the cooling channels 220 may each extend to one or more outlets defined by one or more through holes positioned along the trailing edge of the airfoil 204.
    The bucket 200 may include one or more feed lines 248 (shown by hidden lines) defined in the bucket 200. As shown in FIG. 6, the one or more feed channels 248 may each extend radially from an inlet 252 defined in the stem 206 at the root end 218 of the bucket 200 to the position near the platform 208. In this way, the one or more feed channels 248 may extend through the shaft 206 and into the airfoil 204, respectively. The one or more feed channels 248 may each be in direct communication with one or more of the cooling channels 220 at the joint 254 positioned in the airfoil 204. In particular, the one or more feed channels 248 may be in direct communication with one or more of the radially inner portions 132 at the joint 254. In some aspects, the one or more feed channels 248 may have a fourth cross-sectional area that is constant along the radial length of the feed channel 248. In such aspects, the fourth cross-sectional area may be greater than the first cross-sectional area of each of the one or more radially inner portions 132. Other arrangements may be used in which one or more of the cooling channels 120 may each include additional spaced intermediate portions positioned between the radially inner portion 132 and the radially outer portion 134, each additional spaced intermediate portion has specific and different cross-sectional area.
    FIG. 7 illustrates a cross-sectional top view of a portion of the turbine blade 100 along the outer surface 122 of the airfoil 104, illustrating the structure of one of the radially inner portions 132 of one of the cooling channels 120. The airfoil 104 includes a base portion 160 and a cover layer 164 extending over the base portion 160. In this manner, the cover layer 164 forms the outer surface 122 of the airfoil 104. As shown, the cooling passage 120 is defined by the base portion 160 and the cover layer 164 of the airfoil 104 , In particular, the cooling channel 120 is defined by a channel 168 formed in the base section 160, over which the cover layer 164 extends. The trough 168 may have a substantially rectangular shaped cross section, as illustrated, although the trough 168 may alternatively be configured with a cross section of other shapes. Additionally, gutter 168 may include rounded corners. Along the radially inner portion 132 of the cooling passage 120, the groove 168 may have a width w, and a depth d. Further, along the radially inner portion 132 of the cooling channel 120, the cover layer 164 may have a thickness t 1. The thickness t, may remain constant along the radial length of the radially inner portion 132, resulting in a constant distance between the groove 168 and the outer surface 122 of the airfoil 104. Alternatively, the thickness t 1 may vary along the radial length of the radially inner portion 132, resulting in a varying distance between the groove 168 and the outer surface 122 of the airfoil 104. Although the size of the nominal diameter of the cooling passage 120 may be within a wide range of sizes, in certain embodiments, the nominal diameter may preferably be within 0.105 inches (0.010 inches) and 0.30 inches (7.62 mm). In addition, the cooling channels 120 may have a pitch / nominal diameter ratio that is greater than 1 for larger cooling channels 120, and that may be between 3 and 10 for smaller cooling channels 120.
    FIG. 8 illustrates a cross-sectional top view of a portion of the turbine blade 100 along the outer surface 122 of the airfoil 104, illustrating the structure of one of the radially outer portions 134 of one of the cooling channels 120. Along the radially outer portion 134 of the cooling channel 120, the channel 168 may have a width w0 and a depth d0. Further, the cover layer 164 may have a thickness t0 along the radially outer portion 134 of the cooling channel. The thickness t0 may remain constant along the radial length of the radially outer portion 134, resulting in a constant distance between the groove 168 and the outer surface 122 of the airfoil 104. Alternatively, the thickness t may vary along the radial length of the radially outer portion 134, resulting in a varying distance between the groove 168 and the outer surface 122 of the airfoil 104. As described above, the first cross-sectional area of the radially inner portion 132 of the cooling passage 120 is larger than the second cross-sectional area of the radially outer portion 134 of the cooling passage 120. In addition, the thickness t, of the radially inner portion 132 of the cooling passage 120 may be greater than the thickness to radially outer portion 134 of the cooling channel 120 be. As noted above, although the size of the nominal diameter of the cooling passage 120 may be in a wide range of sizes, in certain embodiments, the nominal diameter may preferably be within 0.010 inches and 0.30 inches. lie. In addition, the cooling channels 120 may have a pitch / nominal diameter ratio that is greater than 1 for larger cooling channels 120, and that may be between 3 and 10 for smaller cooling channels 120.
    The cooling channels 120 of the turbine blade 100 may be formed by a variety of methods. In certain aspects, the channel 168 of the cooling channel 120 may be formed in the base portion 160 of the airfoil 104 by milling, wire-EDM, milled EDM, plunge EDM, water jet cutting, laser cutting, or casting may be used to form the channel 168. After the formation of the channel 168, the cover layer 164 may be formed over the channel 168 in a manner that causes the cover layer 164 to closely adhere to the base section 160. In some aspects, the cover layer 164 may include a thin foil or sheet brazed or welded to the base portion 160. In further aspects, the cover layer 164 may include a spray coating that bridges the gutter 168 and joins the base portion 160. Other methods of forming the cover layer 164 may be In certain aspects, the outlet 126, 128 of the cooling channel s 120 are formed by drilling, water jet cutting or laser cutting. Other methods of forming the outlet 126, 128 may be used. The feed channels 148 of the turbine blade 100 may also be formed by a variety of methods. In certain aspects, the feed channels 148 may be formed in the shank 106 and platform 108 by drilling, STEM drilling, milling, wire EDM, mill EDM, die sinking, water jet cutting, laser cutting, or casting. Other methods of forming the feed channels 148 may be used. Any combination of the above methods may be used to create different cooling channels 120, outlets 126, 128, and feed channels 148.
    During operation of the gas turbine engine 100 containing the turbine blade 100, a cooling fluid, such as a coolant, may be used. Bleed air from the compressor 15 are fed into the feed channels 148 via the inlets 152. The cooling fluid can then pass through the cooling channels 120 and leave the blade 100 via the outlets 126, 128. As a result, heat from the blade 100, and more particularly from the airfoil 104, may transfer to the cooling fluid as it passes through the cooling channels 120 and is then directed into the hot gas path 54 at the tip end 114 of the blade 100. Within the cooling channels 120, the cooling fluid may pass through the radially inner portions 132 and the radially outer portions 134, allowing for transfer of heat from the blade 100 to the cooling fluid at different rates.
    The radially inner portions 132 and the radially outer portions 134 of the cooling holes 120 may be configured to optimize heat transfer from the blade 100, and more particularly from the airfoil 104, to the cooling fluid. Since the cooling requirements toward tip end 114 of blade 100 are greater, the first cross-sectional area of each of radially inner portions 132 is greater than the second cross-sectional area of each of radially outer portions 134 the radially inner portions 132 are minimized. Likewise, the fourth cross-sectional area of each of the delivery channels 148 may be larger than the first cross-sectional area of each of the radially inner portions 132 to minimize heat transfer between the cooling fluid and the blade 100 along the delivery channels 148. Additionally, the cover layer 164 may have a first thickness t over the radially inner portions 132 and a second thickness t0 above the radially outer portions 134. The first thickness t 1 may be greater than the second thickness t to further minimize heat transfer between the cooling fluid and the blade 100 along the radially inner portions 132. As noted above, one or more of the first thickness t, and the second thickness t0 may vary along the radial length of the corresponding radially inner portion 132 or the radially outer portion 134.
    The turbine blade 100 described herein thus provides an improved cooling arrangement to withstand the high temperatures along the hot gas path 54 of the gas turbine engine 10. The bucket 100 includes a number of cooling channels 120, each having a cross-sectional area that varies along the radial length of the cooling channel 120. In particular, one or more of the cooling channels 120 may have one or more radially inner portions 132 having a first cross-sectional area and one or more radially outer portions 134 having a second cross-sectional area configured such that the first cross-sectional area is greater than the second cross-sectional area. In this way, the cooling channels 120 may accommodate the variation in cooling requirements along the radial length of the blade 100 for efficient cooling. Alternatively, one or more of the cooling channels 120 may extend axially along the outer surface of the airfoil 104, may axially reverse, or may form a serpentine pattern or any other pattern for optimizing cooling of the blade 100. As a result, the cooling channels 120 may reduce the amount of airflow needed to cool the blade 100 while increasing the overall efficiency of the gas turbine engine 10. Further, the cooling channels 120 of the blade 100 may allow for various complex three-dimensional shapes or rotation since the cooling channels 120 may be formed along the outer surface 122 of the airfoil 104. In this way, the airfoil 104 may be configured for improved aerodynamics without regard to including cooling holes with a straight track.
    Although the embodiments presented above have been illustrated and described with reference to a blade 100, 200, it will be understood that similar arrangements of a cooling channel 120, 220 may be used on or along the hot gas path of the gas turbine 10, such as eg with a vane or a shroud. Further, although the embodiments presented above are illustrated and described with cooling channels 120, 220 extending radially inwardly of the blade 100, 200, the cooling channels 120, 220 may extend axially along the blade, it should be understood that they are axially reversible may or may form a serpentine pattern or any other pattern or combination thereof to optimize cooling of the blades 100, 200.
    The present application and the resulting patent provide a turbine blade for a gas turbine. The turbine blade may include a platform, an airfoil extending radially out of the platform, and a number of cooling channels defined in the airfoil and in the vicinity of an outer surface of the airfoil. Each of the cooling channels includes a radially inner portion having a first cross-sectional area and at least one radially outer portion having a second cross-sectional area, wherein the first cross-sectional area is larger than the second cross-sectional area.
    10 gas turbine 15 compressor 20 airflow 25 burners 30 fuel flow 35 combustion gas flow 40 turbine 45 shaft 50 external load 52 turbine stages 54 hot gas path 56 first stage 58 first stage vanes 60 first stage buckets 62 second stage 64 second stage vanes 66 buckets second stage 68 third stage 70 third stage vanes 72 third stage vanes 76 airfoil 78 shank 80 platform 82 tip shroud 84 tip end 86 foot end 88 cooling holes 90 outlets 92 feed holes 94 inlets 96 joint 100 turbine blade 104 airfoil 106 shank 108 platform 112 tip shroud 114 tip end 118 foot end 120 cooling channels 122 outer surface 126 outlets 128 outlets 132 radially inner section 134 radially outer section 136 connection point 138 radially intermediate section 142 connection point 144 connection point 148 supply channels 152 inlet 1 54 Junction 160 Base 164 Top 168 Gutter 200 Turbine Blade 204 Airfoil 206 Shaft 208 Platform 212 Tip Shroud 214 Tip End 218 Foot End 220 Cooling Channels 222 Outer Surface 226 Outlet 228 Outlet 132 Radial Inner Section 134 Radial Outer Section 236 Junction 238 Radially Intermediate Section 242 Junction 244 Junction 248 Feed Channels 252 inlet 254 connection point
    权利要求:
    Claims (8)
    [1]
    claims
    A turbine blade (100, 200) for a gas turbine, the turbine blade (100, 200) comprising: a platform (108, 208); an airfoil (76, 104, 204) extending radially from the platform (108, 208); and a number of cooling passages (120, 220) defined in the airfoil (76, 104, 204) adjacent the outer surface (122, 222) of the airfoil (76, 104, 204), each of the cooling passages (120, 220) having a radially inner portion (132). with a first cross-sectional area defined by the cooling passages (120, 220) and at least one radially outer portion (134) having a second cross-sectional area defined by branched cooling passages (120, 220), and wherein the first cross-sectional area is greater than the second cross-sectional area, wherein the airfoil ( 76, 104, 204) has a base portion (160) and a cover layer (164) extending above the base portion (160), each of the cooling channels (120, 220) being defined by a groove (168) formed in the base portion (160) is formed and over which the cover layer (164) extends.
    [2]
    The turbine blade (100, 200) of claim 1, wherein at least one of the cooling channels (120, 220) has a radially intermediate portion (138) extending between the radially inner portion (132) and at least one radially outer portion (134 ), wherein the radially intermediate portion (138) has a third cross sectional area, and wherein the third cross sectional area is larger than the second cross sectional area and smaller than the first cross sectional area.
    [3]
    The turbine blade (100, 200) of claim 1, wherein the cover layer (164) has a first thickness above the radially inner portion (132) and a second thickness above the radially outer portion (134), and wherein the first thickness is greater than the second thickness is.
    [4]
    The turbine bucket (100, 200) of claim 1, wherein at least one of the cooling channels (120, 220) extends radially along the outer surface (122, 222) of the airfoil (76, 104, 204) from the platform (108, 208) to one or more of a plurality of outlets (126, 128, 226) defined in a tip shroud (82) of the turbine bucket (100, 200); and / or wherein at least one of the cooling channels (120, 220) extends radially along the outer surface (122, 222) of the airfoil (76, 104, 204) from the platform (108, 208) to one or more in the outer surface ( 122, 222) of the airfoil (76, 104, 204) and outlets (126, 128, 226) positioned between the platform (108, 208) and a tip shroud (82) of the turbine bucket (100, 200).
    [5]
    The turbine blade (100, 200) of claim 1, further comprising a shaft (78) extending radially from said platform (108, 208) and away from said airfoil (76, 104, 204) and at least one in said shaft (10). 78), the supply channel (148, 248) communicating with one or more of the cooling channels (120, 220) at a junction (154, 254).
    [6]
    The turbine blade (100, 200) of claim 5, wherein the supply channel (148, 248) communicates with one or more of the radially inner portions (132) at the joint (254), and wherein the joint (254) in the airfoil (76, 104, 204) is positioned between the platform (108, 208) and a tip shroud (82) of the turbine blade (100, 200), and wherein the feed channel (148, 248) has a fourth cross sectional area, the fourth Cross-sectional area is greater than the first cross-sectional area.
    [7]
    7. hot gas path component of a gas turbine (10) with a turbine blade (100, 200) according to any one of the preceding claims.
    [8]
    The hot gas path component of claim 7, wherein the cooling passages (120, 220) extend radially along the hot gas path component, a first portion having the radially inner portions (132) and at least a second portion having at least one radially outer portion (134), and wherein the cooling channels (120, 220) extend axially along the hot gas path component, the first portion being axially downstream of at least a second portion, and wherein the hot gas path component comprises a vane (58, 64, 70) and the cooling channels (120, 220 ) extend along the vane (58, 64, 70).
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    同族专利:
    公开号 | 公开日
    CH707844A2|2014-09-15|
    JP2014177943A|2014-09-25|
    DE102014103007A1|2014-09-18|
    CN203835465U|2014-09-17|
    US9567859B2|2017-02-14|
    JP6438662B2|2018-12-19|
    US20140286771A1|2014-09-25|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    优先权:
    申请号 | 申请日 | 专利标题
    US13/804,159|US9567859B2|2013-03-14|2013-03-14|Cooling passages for turbine buckets of a gas turbine engine|
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