• 专利附录
  • 专利说明
  • 权利要求
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    • 专利摘要:
    The present invention provides a stationary blade cooling structure including: an end wall (204) connected to a radial end of an airfoil (150) relative to a rotor axis of a turbomachine; and a substantially crescent-shaped chamber (218) positioned within the end wall (204) and radially offset from a trailing edge (154) of the airfoil (150), the substantially crescent-shaped chamber (218) containing a cooling fluid from a cooling circuit (216 ), wherein the substantially crescent-shaped chamber (218) extends from a front portion (222) positioned near one of a pressure side surface (156) and a suction side surface (158) of the airfoil (150) to a rearward one Portion (154) positioned near the trailing edge (154) of the airfoil (150) and the other of the pressure side surface (156) and the suction side surface (158) of the airfoil (150), the rear region (224) the substantially crescent-shaped chamber (218) is in fluid communication with the front portion (222) of the substantially crescent-shaped chamber (218).
    公开号:CH711387A2
    申请号:CH00882/16
    申请日:2016-07-11
    公开日:2017-01-31
    发明作者:Christopher Lee Golden；Michael Earnhardt Dustin；Jessica Iduate Michelle；Bryan David Lewis；Christopher Donald Porter；Wayne Weber David
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    BACKGROUND OF THE INVENTION
    The disclosure generally relates to stationary blades, and more particularly to a stationary blade cooling structure.
    Stationary blades are used in turbine applications to direct hot gas flows to moving blades to produce power. In steam and gas turbine applications, the stationary vanes are referred to as vanes and are mounted to an outer structure, such as a housing, and / or an inner seal structure by end walls. Each end wall is connected to a corresponding end of a stationary blade airfoil. Stationary blades may also include passages or other means for circulating cooling fluids that absorb heat from operating components of the turbomachine.
    In order to work in facilities with extreme temperatures, the airfoil and the end walls must be cooled. For example, in some devices, cooling fluid is withdrawn from the impeller space and directed to inner stationary walls of the stationary blade for cooling. On the other hand, in many gas turbine applications, later stage vanes may be provided with a cooling fluid, e.g. be fed with air, which is taken from the compressor. Outer peripheral end walls may receive the cooling fluid directly, while the inner peripheral end walls may receive the cooling fluid after it has been passed through the airfoil from the outer periphery. In addition to the effectiveness of cooling, the structure of a stationary blade and its components may affect other factors such as manufacturability, ease of testing, and durability of a turbomachine.
    BRIEF DESCRIPTION OF THE INVENTION
    [0004] A first aspect of the present disclosure provides a stationary blade cooling structure including: an end wall connected to a radial end of an airfoil relative to a rotor axis of a turbomachine, the airfoil having a pressure side surface, a suction side surface Contains leading edge and a trailing edge; and a substantially crescent-shaped chamber positioned within the end wall and radially offset from the trailing edge of the airfoil, wherein the substantially crescent-shaped chamber receives a cooling fluid from a refrigeration circuit, the substantially crescent-shaped chamber extending from a forward region formed in the airfoil Near one of the pressure side surface and the suction side surface of the airfoil is arranged, extends to a rear portion which is disposed in the vicinity of the trailing edge of the airfoil and the other of the pressure side surface and the suction side surface of the airfoil, wherein the cooling fluid in the front region with a is in heat communication with the pressure side surface and the suction side surface of the airfoil, wherein the cooling fluid in the rear region is in thermal communication with a portion of the end wall in the vicinity of the trailing edge of the airfoil and wherein the rear region of the substantially L-shaped chamber is in fluid communication with the front portion of the substantially sickle-shaped chamber.
    In the aforementioned cooling structure, the rear portion of the substantially crescent-shaped chamber may extend substantially perpendicularly from one end of the front portion of the substantially crescent-shaped chamber.
    Any of the above-mentioned cooling structures may further include a plurality of heat-conducting means extending through at least one of the front portion and the rear portion of the substantially crescent-shaped chamber.
    In addition, the substantially crescent-shaped chamber may further include a perimeter wall and further include a plurality of access areas positioned substantially along the perimeter wall of the substantially crescent-shaped chamber, each of the plurality of access areas being free of heat conducting means therein.
    Further additionally or as an alternative, the substantially crescent-shaped chamber may further include a transition region substantially radially aligned with the trailing edge of the airfoil and disposed between the forward region and the rearward region of the substantially crescent-shaped chamber, the transition region is free of heat conducting facilities in it.
    In the cooling structure of any of the above-mentioned types, an axial length component of the front portion of the substantially crescent-shaped chamber may be at least about half of an axial length component of one of the pressure side surface and the suction side surface of the airfoil.
    In some embodiments of any of the aforementioned cooling structures, the substantially crescent-shaped chamber may further include a transition region that is substantially radially aligned with the trailing edge of the airfoil and disposed between the forward region and the rearward region of the substantially crescent-shaped chamber, and further comprising a protrusion extending from a radial surface of the transition region, the protrusion configured to direct the cooling fluid from the forward region to the rearward region of the substantially sickle-shaped chamber.
    In the last mentioned embodiments, a width of the transition region between the front portion and the rear portion may correspond to approximately an axial width of the trailing edge of the airfoil.
    In any of the aforementioned cooling structures, the substantially crescent-shaped chamber may comprise one of at least two generally crescent-shaped chambers positioned within the end wall, the airfoil being one of a pair of airfoils extending from the end wall protrude substantially radially.
    A second aspect of the present disclosure provides a stationary blade that includes: an airfoil that includes a pressure side surface, a suction side surface, a leading edge, and a trailing edge, the airfoil further including a cooling circuit therein; an end wall connected to a radial end of the airfoil relative to a rotor axis of a turbomachine; and a substantially crescent-shaped chamber positioned within the end wall and radially offset from the trailing edge of the airfoil, the substantially crescent-shaped chamber receiving a cooling fluid from the cooling circuit, the substantially crescent-shaped chamber extending from a forward region formed in the airfoil Near one of the pressure side surface and the suction side surface of the airfoil is arranged, extends to a rear portion which is disposed in the vicinity of the trailing edge of the airfoil and the other of the pressure side surface and the suction side surface of the airfoil, wherein the cooling fluid in the front region one of the pressure side surface and the suction side surface of the airfoil is in thermal communication, wherein the cooling fluid is in the rear region with a portion of the end wall in the vicinity of the trailing edge of the airfoil in thermal communication and wherein the rear portion of sic essentially sic helical chamber is in fluid communication with the front portion of the substantially sickle-shaped chamber.
    In the aforementioned stationary blade, the rear portion of the substantially crescent-shaped chamber may extend substantially perpendicularly from the front portion of the substantially crescent-shaped chamber.
    In some embodiments, any stationary blade mentioned above may further include a plurality of heat-conducting devices extending through at least one of the front portion and the rear portion of the substantially crescent-shaped chamber.
    In the last-mentioned embodiments, the substantially crescent-shaped chamber may further include a transition region substantially radially aligned with the trailing edge of the airfoil and disposed between the forward region and the rearward region of the substantially crescent-shaped chamber, the transition region being exposed of thermally conductive means therein.
    In addition, a width of the transition region between the front portion and the rear portion may be approximately equal to an axial width of the trailing edge of the airfoil.
    Additionally or alternatively, the substantially crescent-shaped chamber may further include a perimeter wall and further include a plurality of access areas substantially along the perimeter wall of the substantially crescent-shaped chamber, each of the plurality of access areas being free of heat conducting means therein ,
    [0019] In some embodiments of any stationary blade mentioned above, the substantially crescent-shaped chamber may further include a transition region substantially radially aligned with the trailing edge of the airfoil and disposed between the forward region and the rearward region of the substantially crescent-shaped chamber, and may further include a protrusion extending from a radial surface of the transition region, wherein the protrusion is configured to direct the cooling fluid from the forward region to the rearward region of the substantially sickle-shaped chamber.
    A third aspect of the present disclosure provides a turbine nozzle pair including: a first airfoil having a first cooling circuit therein; an end wall connected to a radial end of the first airfoil relative to a rotor axis of a turbomachine; a second airfoil having a second cooling circuit therein, wherein the second airfoil is aligned substantially parallel to the first airfoil, the endwall being connected to a radial end of the airfoil relative to the rotor axis of the turbomachine, and each of the first airfoil and the first airfoil the second airfoil further includes a pressure side surface, a suction side surface, a leading edge, and a trailing edge; a first substantially crescent-shaped chamber positioned within the end wall and radially offset from the trailing edge of the first airfoil, the first substantially crescent-shaped chamber receiving a first cooling fluid from the first cooling circuit, the first substantially crescent-shaped chamber extending from a forward A region disposed near one of the pressure side surface and the suction side surface of the first airfoil extends to a rear region located near the trailing edge of the first airfoil and the other of the pressure side surface and the suction side surface of the first airfoil; wherein the first cooling fluid in the front portion of the first substantially crescent-shaped chamber is in thermal communication with one of the pressure side surface and the suction side surface of the first airfoil, wherein the first cooling fluid is substantially in the rear portion of the first -shaped chamber with a portion of the end wall is in the vicinity of the trailing edge of the first blade body in thermal communication, and wherein the rear portion of the first substantially crescent shaped chamber with the front portion of the first substantially crescent shaped chamber in fluid communication; and a second substantially crescent-shaped chamber positioned within the end wall and radially offset from the trailing edge of the second airfoil, the second substantially crescent-shaped chamber receiving a second cooling fluid from the second cooling circuit, the second substantially crescent-shaped chamber extending from a second front portion, which is disposed in the vicinity of one of the pressure side surface and the suction side surface of the second airfoil, to a rear portion disposed in the vicinity of the trailing edge of the second airfoil and the other of the pressure side surface and the suction side surface of the second airfoil wherein the second cooling fluid in the front portion of the second substantially crescent-shaped chamber is in thermal communication with one of the pressure side surface and the suction side surface of the second airfoil, wherein the second cooling fluid is in the rear portion of the second in the Wes The sickle-shaped chamber is in thermal communication with a portion of the end wall near the trailing edge of the second airfoil and wherein the rear portion of the second substantially crescent-shaped chamber is in fluid communication with the front portion of the second substantially crescent-shaped chamber.
    The aforementioned turbine nozzle pair may further include a plurality of heat conducting means extending through one of the first and second substantially crescent-shaped chambers in one of the front portion and the rear portion thereof.
    Further, one of the first and second substantially crescent-shaped chambers may further include a transition region disposed between the front portion and the rear portion thereof and substantially radially aligned with the trailing edge of one of the first airfoil and the second airfoil is, wherein the transition region is free of thermally conductive means therein.
    Still further, the turbine nozzle pair may further include a protrusion extending from an axial surface of the transition region, the protrusion configured to direct the cooling fluid from the forward region to the rearward region of the substantially sickle shaped chamber.
    BRIEF DESCRIPTION OF THE DRAWINGS
    These and other features of this invention will become more readily apparent from the following detailed description of various embodiments of the invention, taken in conjunction with the accompanying drawings, which illustrate various embodiments of the invention, wherein:<Tb> FIG. 1 <SEP> shows a schematic view of a turbomachine.<Tb> FIG. 2 <SEP> shows a cross-sectional view of a stationary blade airfoil positioned in a flow path of an operating fluid according to embodiments of the present disclosure.<Tb> FIG. Figure 3 shows a cross-sectional view of a stationary blade between two rotor blades in a turbine section of a turbomachine.<Tb> FIG. 4 <SEP> is a cutaway perspective view of a stationary blade cooling structure according to embodiments of the present disclosure.<Tb> FIG. FIG. 5 shows a cutaway partial perspective view of a chamber within an end wall in accordance with embodiments of the present disclosure. FIG.<Tb> FIG. Figure 6 provides a cut away partial perspective view of a transition area of a chamber according to embodiments of the present disclosure.<Tb> FIG. FIG. 7 shows an enlarged partial cutaway perspective view of a transition area of a chamber according to embodiments of the present disclosure. FIG.
    It should be noted that the drawings of the invention are not necessarily to scale. The drawings are intended to depict only typical aspects of the invention and, thus, should not be considered as limiting the scope of the invention. In the drawings, like reference numerals represent like elements between the drawings.
    DETAILED DESCRIPTION OF THE INVENTION
    Embodiments of the present disclosure generally relate to stationary blade cooling structures. In particular, embodiments of the present disclosure provide an end wall connected to a radial end of a stationary blade airfoil, the airfoil including a pressure side surface, a suction side surface, a leading edge, and a trailing edge. The end wall may include a substantially crescent-shaped chamber therein which is radially offset from the airfoil. The substantially crescent-shaped chamber may, as further described herein, be radially offset therefrom and positioned proximate the pressure side surface, the trailing edge, and / or the suction side surface such that the chamber partially engages an outer contour of the airfoil. During operation, cooling fluids may enter and pass through the chamber to absorb heat from portions of the end wall that are proximate the respective surfaces of the airfoil (i.e., the pressure side surface, the trailing edge, and / or the suction side surface). The chamber may include a front portion positioned near the pressure side surface or the suction side surface of the airfoil and a rear portion positioned proximate to at least the trailing edge of the airfoil and the other from the pressure side surface or the suction side surface of the airfoil , The front portion and the rear portion may together form separate portions of the crescent geometry of the chamber.
    Spatially relative expressions, such as "inner," "outer," "below," "below," "lower," "above," "upper," "inlet," "outlet," and the like, can may be used herein for ease of description to describe the relationship of an element or device to one or more other element (s) or devices as illustrated in the figures. Spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientation shown in the figures. For example, if the device in the figures were turned over, the elements described herein as "below" or below "other elements or devices would then be located" above "the other elements or devices. Thus, the exemplary term "under" may include both a positioning above and below. The device may be otherwise oriented (rotated 90 ° or to other orientations) and the spatially relative descriptions as used herein interpreted accordingly.
    As mentioned above, the disclosure provides a cooling structure for a stationary blade of a turbomachine. In an embodiment, the cooling structure may include a substantially crescent-shaped chamber having a front portion near a pressure side surface or a suction side surface of an airfoil. The substantially crescent-shaped chamber may extend from the front portion to a rear portion near a trailing edge and the opposite pressure side surface or suction side surface of the airfoil. 1 shows a turbomachine 100 that includes a compressor section 102 that is operatively connected to a turbine section 104 via a common compressor / turbine shaft 106. The compressor section 102 is further fluidly connected to the turbine section 104 via a combustor assembly 108. The combustor assembly 108 includes one or more combustors 110. The combustors 110 may be mounted to the turbomachine 100 in a wide variety of configurations including, but not limited to, an annular tubular arrangement. The compressor section 102 includes a plurality of compressor impellers 112. The impellers 112 include a first stage compressor impeller 114 having a plurality of first stage compressor impellers 116, each having an associated airfoil portion 118. Likewise, the turbine section 104 includes a plurality of turbine runners 120 that include a first stage turbine runner 122 having a plurality of first stage turbine runners 124. According to an exemplary embodiment, a stationary blade 200 (FIG. 3) having a cooling structure according to embodiments of the present disclosure may include cooling at end walls and airfoils, e.g. are arranged in the turbine section 104, achieve. However, it is understood that embodiments of the stationary blade 200 and the various cooling structures described herein may be positioned in other components or regions of the turbomachine 100.
    Referring to Figure 2, a cross section of a flow path 130 for operating fluids containing an airfoil 150 therein is illustrated. The airfoil 150 may be part of the stationary blade 200 (FIG. 3) and may further include the components and / or datums as described herein. The positions on the airfoil 150, identified in FIG. 2 and discussed herein, are provided as examples and are not intended to limit possible positions and / or geometries for the airfoils 150 in accordance with embodiments of the present disclosure. The placement, setup and alignment of various subcomponents may vary depending on the intended use and the type of power generation system in which cooling structures according to the present disclosure are used. The shape, curvatures, lengths, and / or other geometric features of the airfoil 150 may also vary based on the application of a particular turbomachine 100 (FIG. 1). The airfoil 150 may be positioned between successive turbine blades 124 (FIG. 1) of a power generation system, such as the turbomachine 100.
    The airfoil 150 may be positioned downstream of a turbine blade 124 (FIG. 1) and upstream of another subsequent turbine blade 124 (FIG. 1) in a flow path for an operating fluid. Fluids may pass over the airfoil 150, e.g. along one or more paths F, as they flow from one turbine blade 124 to another. A leading edge 152 of the airfoil 150 may be positioned at an initial contact point between the operating fluid in the flow path 130 and the airfoil 150. On the other hand, a trailing edge 154 may be positioned on the opposite side of the airfoil 150. In addition, the airfoil 150 may include a pressure side surface 156 and / or a suction side surface 158 that are distinguished by a transverse line that bisects the leading edge 152 substantially and extends to the apex of the trailing edge 154. The pressure side surface 156 and the suction side surface 158 may also be distinguished from one another based on whether fluids in the flow path 130 exert a positive or negative resultant pressure against the airfoil 150. A portion of the pressure side surface 156 positioned proximate the trailing edge 154 may be known and designated as a "high Mach number region" of the airfoil 150, based on the fluids in that region as compared to other surfaces of the airfoil Flow airfoil 150 at a higher speed.
    Referring to FIG. 3, a cross section of a flow path 130 is illustrated past a stationary blade 200 positioned within the turbine section 104. Operating fluid (e.g., hot combustion gases, steam, etc.) may flow through the flow path 130 (e.g., along flow lines F) to reach other turbine blades 124 as directed by the position and contours of the stationary blade 200. The turbine section 104 is illustrated as extending along a rotor axis Z of the turbine runner 122 (e.g., coaxial with the shaft 106 (Figure 1)) and with a radial axis R extending outwardly therefrom. The stationary blade 200 may include an airfoil 150 which is oriented substantially along the radial axis R (i.e., extending in a direction substantially parallel thereto, i.e., within about 10 degrees of the same angle plane). Although a single stationary blade 200 is illustrated in the cross-sectional view of FIG. 3, it will be understood that multiple turbine blades 124 and stationary blades 200 may extend radially from the turbine runner 122, e.g. can extend laterally into and / or out of the plane of the page. An airfoil 150 of the stationary blade 200 may include two end walls 204, one connected to an inner radial end of the airfoil 150, and another connected to an opposite outer radial end of the airfoil 150.
    An end wall 204 may be positioned proximate to the turbine runner 122, which is disposed substantially on an inner radial surface, while another end wall 204 may be positioned proximate a turbine shell 212 that is substantially on an outer radial surface is arranged. During operation, the hot combustion gases that propagate along the flow lines F may apply heat to the airfoil 150 and / or the end wall or end walls 204, e.g. by means of effective fluids which come into contact with the airfoil 150 and the end wall or end walls 204 of the stationary blade 200. The airfoil 150 of the stationary blade 200 may include a cooling circuit 216 therein. The cooling circuit 216 may include or be provided as a cavity within the airfoil 150 within the airfoil 150 for transmitting cooling fluids radially through the airfoil 150, the cooling fluids receiving heat from the active fluid in the flow path 130 via the heat-conductive material composition of the airfoil 150 can absorb.
    The cooling circuit 216, which may be in the form of a baffle cavity, may circulate a cooling fluid through a partially hollow interior of the airfoil 150 between the two end walls 204. An impingement refrigeration cycle generally refers to a refrigeration cycle structured to provide a cooling fluid film at a portion of a cooled component (eg, a transverse radial member of the airfoil 150), thereby transferring the thermal energy of substances outside the cooled component to an internal volume of the air cooled component is reduced. Cooling fluids in the cooling circuit 216 may originate from a chamber 218 and / or flow to a chamber 218 positioned within a single end wall 204 or within both end walls 204. Cooling fluids in the chamber (s) 218 that have not flowed through the cooling circuit 216 may be referred to as "pre-crash" cooling fluids, while cooling fluids in the chamber (s) 218 previously passed through the cooling circuit 216 have flowed, can be referred to as "rebound" cooling fluids. Embodiments of the present disclosure may include, but are not limited to, a stationary blade cooling structure 200 having a chamber 218 for absorbing heat from a plurality of surfaces of the end wall or end walls 204 proximate the location where the airfoil 150 has the end wall (s) 204, provide.
    Referring to Figure 4, a cut away perspective view of an end wall 204 having two chambers 218 near a cross section of the two airfoils 150 is illustrated. Each airfoil 150 may extend radially from end wall 204, i. projecting substantially perpendicularly relative to the rotor axis of the turbomachine 100 (Figure 1). As used herein, the term "substantially perpendicular" or "substantially perpendicular" refers to an angle of 90 degrees or an angle that differs from 90 degrees by an insubstantial size, e.g. is within a range of between about 85 degrees and about 95 degrees. Although illustrated in FIG. 4 as an example are two airfoils 150 connected to the end wall 204 (ie, in a turbine vane pair configuration), it will be understood that any desired number of airfoils 150 may be connected to the end wall 204 to suit various turbomachinery designs and applications. Each airfoil 150 may have any of a variety of airfoil constructions and / or implementations, and as an example, may be one of a self-supporting turbine nozzle airfoil blades 150 and / or a second stage airfoil of the turbomachine 100. Likewise, the end wall 2014 may include two chambers 218, each corresponding to a single airfoil 150 in a pair or doublet configuration, or any desired number of chambers 218 therein to suit various applications.
    One or more inlets 220 may be provided for fluid communication between each chamber 218 and a source of cooling fluid, e.g. a cooling circuit or multiple cooling circuits 216, provide. Each chamber 218 may be substantially crescent shaped. As used herein, the term "substantially sickle-shaped" may include any geometry that includes two branching independent paths that originate from the same point of convergence and extend in at least one common direction. As examples, a sickle shape according to this definition may include a C shape, a V shape, a J shape, a bow, a boomerang shape, a crescent shape or an armpit shape, etc. Regardless of the type of generally crescent shape, one end of the chamber 218 may be positioned near the pressure side surface 156 or the suction side surface 158 of the airfoil 150, while an opposite end of the chamber 218 may be proximate the opposing pressure or suction side surface 156, 158 of the airfoil 150 may be positioned. The chamber 218 may thus extend around or below the trailing edge 154 of the airfoil 150. In addition, two portions of the chamber 218 may converge radially below the trailing edge 154 of the airfoil 150. The generally crescent-shaped geometry of the chamber 218 may thus provide an encompassing geometry that substantially follows the contours of the airfoil 150 along portions of the pressure side surface 156 and / or the suction side surface 850, but extends radially below the trailing edge 154.
    Each chamber 218 may include a front portion 222 and a rear portion 224 therein. The front portion 222 may be positioned near the pressure side surface 156 or the suction side surface 158, i. only be separated by the material composition of the end wall 204 of this. The front portion 222 is illustrated in FIG. 4 as an example as it is near the pressure side surface 156, but in alternative embodiments may be near the suction side surface 158. In addition, as illustrated in FIG. 4, the forward portion 222 of the chamber 218 may be positioned proximate the pressure side surface 156 of a corresponding airfoil 150, while also located proximate the suction side surface 158 of another airfoil 150. The rear region 224 may be positioned proximate to both the trailing edge 154 and the opposed pressure side surface 156 or suction side surface 158 relative to the front region 222. The front portion 222 and the rear portion 224 may be distinguishable only based on their position relative to surfaces of the airfoil 150, but it will be understood that other structural features, such as an additional portion or structure, may be used the front portion 222 and the rear portion 224, as otherwise explained herein, may further distinguish the front portion 222 of the chamber 218 from the rear portion 224 of the chamber 218.
    During operation of the turbomachine 100 (FIG. 1), cooling fluids may enter the chamber 218 through the inlet (s) 220 to subsequently pass through the forward region 222 and the rearward region 224 before passing through the chamber 218 leave the outlet or outlets 226. Each chamber 218 may have inlets of e.g. a single cooling circuit 216 or multiple cooling circuits 216 of the respective airfoils 150 included. Cooling fluids in the forward portion 222 of the chamber 218 may absorb heat from a portion of the end wall 204 that is positioned proximate the pressure side surface 156 or the suction side surface 158 of the airfoil 150 as it passes therethrough, e.g. via heat transfer from the airfoil 150 to the chamber 218 through the end wall 204. Cooling fluids in the rear portion 224 of the chamber 218 may absorb heat from a portion of the end wall 204 positioned proximate the pressure side surface 156 or the suction side surface 158 (opposite the surface proximate the front portion 222) and the trailing edge 154 of the airfoil 150 is absorb as they pass through it. The front portion 222 and rear portion 224 of the chamber 218 may converge radially below the trailing edge 154 of the airfoil 150. In other embodiments, as discussed in further detail herein, the forward region 222 and the rearward region 224 may converge at a transition region 236 (FIG. spans the same axial length as the trailing edge 154 of the airfoil 150. The rear region 224 and the forward region 222 may extend from one another and / or from the transition region 236 substantially perpendicularly relative to each other and within the same radial plane of the end wall 204.
    The front portion 222 and the rear portion 224 of the chamber 218 may be configured to have different dimensions and / or contours. In one embodiment, the forward region 222 may have an axial length (e.g., along the axis Z) that measures at least about half the axial length of the airfoil 150 along the proximate pressure side surface 156 or suction side surface 158. In contrast, the rear portion 224 may extend over less than half an axial length of the opposed pressure side surface 156 or suction side surface 158 of the airfoil 150. The axial length of the rear portion 224, which is smaller than an axial length of the front portion 222, may cause the front portion 222 to be significantly larger than the rear portion 224 such that the substantially crescent-shaped chamber 218 has a J-shape having.
    Together, referring to FIGS. 4 and 5, embodiments of the present disclosure may include any number of thermally conductive devices ("devices") 230, such as a socket, within the chamber (s) 218 (eg, within the forward region 222 or the rear portion 224) to transfer heat from the stationary blade 200 to cooling fluids within the chamber (s) 218. In particular, each device 230 may transfer heat from the end wall 204 to cooling fluids therein by increasing the contact area between the cooling fluids that flow through the chamber (s) 218 and the material composition of the end wall 204. The means 230 may be provided as any conceivable means for increasing the contact area between cooling fluids and heat-conducting surfaces, and may, for example, be in the form of shoulders, recesses, projections, cones, walls and / or other devices of other shapes and sizes. In addition, the devices 230 may take a variety of forms, including those with cylindrical geometries, substantially pyramidal geometries, irregular geometries with four or more surfaces, etc. In any event, one or more devices 230 may be positioned within the chamber 218 at a location of the cooling fluid flow path that is downstream of the inlet 220 and upstream of the outlet or outlets 226.
    The positioning of the devices 230, in addition to improving the heat transfer between the end wall 204 and the cooling fluids therein, may increase the temperature differential between the cooling fluids within the forward region 222 and the rearward region 224. The distance between adjacent devices 230 may be sized to account for inspection and testing by particular instruments. An inspection of the stationary blade 200 may e.g. include contacting a precast component of stationary bucket 200 and / or a partially constructed stationary bucket 200 or end wall 204 with a boroscopic lens or other machine for testing the properties of a material. For example, For example, adjacent devices 230 may have a sufficient separation distance to allow a borescope lens or other part of a test equipment to be placed within the chamber (s) 218 between different sockets 230. The spacing between the sockets may vary between applications, and as an example between e.g. about a millimeter (mm) and about 20 mm to account for a range of borescope diameters. In some embodiments, the sockets 230 in the chamber 218 may be partially or completely missing. The chamber 218 may also be bounded by a containment wall 232 that extends a predetermined radial length of the end wall 204, thereby defining a height dimension of the chamber 218. In embodiments where the chamber 218 includes the pedestals 230 therein, the chamber 218 may also include a plurality of access regions 234 positioned substantially along portions of the perimeter wall 232. Each access area 234 may be free of sockets 230 therein, thereby providing additional space for performing inspections of the chamber 218 with a horoscope and / or other tools.
    Referring to Figure 6, a partial perspective cut away view of the end wall 204 with the chambers 218 therein is illustrated. One or more chambers 218 of the end wall 204 may further include a transition region 236 positioned between the front portion 222 and the rear portion 224 of the chamber 218. To increase the heat transfer rate from the airfoil 150 to cooling fluids in the chamber 218, the transition region 236 may be substantially radially aligned with the trailing edge 154 of the airfoil 150. In addition, the transition region 236 may optionally include means 230 therein for increasing the flow rate of cooling fluids through the transition region 236. In an alternative embodiment, the transition region 236 may be devoid of devices 230 therein. In order to branch off a portion of the cooling fluid that has not absorbed heat in the rearward region 224, one or more outlets 226 may be at least partially in fluid communication with the transition region 236. To further provide heat transfer from the trailing edge 154 of the airfoil 150, an axial width of the transition region 236 between the forward region 222 and the rearward region 224 may correspond approximately to an axial width of the trailing edge 154 such that substantially no devices 230 radially below the trailing edge 154 Trailing edge 254 are positioned.
    Referring to Figure 7, a partial perspective view of the transition region 236 is illustrated in further detail. The transition region 236 may optionally include a protrusion 238 therein, e.g. extends from an upper or lower radial surface of the chamber 218 to direct cooling fluids in the forward region 222 into the rearward region 224 of the chamber 218. The protrusion 238 may be in the form of an elongated device, such as illustrated by way of example in FIG. 7, and may be constructed of the same thermally conductive material as the end wall 204 or of another thermally conductive material. As further illustrated in FIG. 7, the protrusion 238 may have a different shape than the device (s) 230, such as an elongated bulkhead, a swirler, a vane, etc., to direct at least a portion of the cooling fluids into the rear region 224 of the chamber 218 to direct. In operation, the protrusion 238 may be thermally conductive or heat-insulating based on whether further heat absorption in the transition region 238 is desired. The projection 238 may also direct portions of the cooling air within the chamber 218 into the rear portion 224 and / or into other components via the outlets 226 during operation.
    Embodiments of the present disclosure may provide various technical and commercial advantages, some of which are exemplified herein. For example, For example, providing a substantially crescent-shaped chamber within the end wall 204 may improve heat exchange communication between various surfaces of the airfoil 150 and cooling fluids within the end wall 204. Among other things, an improved heat exchange connection can reduce the total amount of vane cooling flow needed during operation, and can reduce the complexity of the design required to end walls 204 from cast ferrous metal substances such as aluminum, copper, iron, lead and / or combinations of these materials. The substantially crescent shape of the chamber 218 with a point of convergence radially offset from the trailing edge 154 of the airfoil 150 can reduce the mechanical rigidity of the chamber 218. This reduction in mechanical stiffness can provide mechanical benefits derived therefrom, such as improved manufacturability and / or durability.
    The apparatus and method according to the present disclosure are not limited to any particular gas turbine engine, combustion engine, any particular power generation system or other system, and may be used with power generation systems and / or systems (e.g., combined cycle power plants, single cycle power plants, nuclear reactors, etc.). Moreover, the device according to the present invention may be used with other systems not described herein which may benefit from the increased operating range, increased efficiency, greater durability and reliability of the device described herein. In addition, the various injection systems may be used together on a single vane or on / with different vanes in different sections of a single power generation system. Any number of different embodiments may be added or shared, if desired, and the embodiments described herein by way of example are not intended to be mutually exclusive.
    The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. As used herein, the singular forms "a," "an," "the," "the," and "the" are also intended to encompass the plural forms unless the context clearly dictates otherwise. It is further understood that the terms "having" and / or "having" when used in this specification specify the presence of the specified features, integers, steps, operations, elements, and / or components, but the presence or inclusion one or more characteristics, integers, steps, operations, elements, components and / or their groups are not mutually exclusive.
    This written description uses examples to disclose the invention, including the best mode, and to enable any person skilled in the art to practice the invention, including the creation and use of any devices or systems and performing any incorporated methods , The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
    Embodiments of the present disclosure provide a stationary blade cooling structure that may include: an end wall 204 connected to a radial end of an airfoil 150 relative to a rotor axis of a turbomachine 100; and a substantially crescent-shaped chamber 218 positioned within the end wall 204 and radially offset to a trailing edge 154 of the airfoil 150, the substantially crescent-shaped chamber 218 receiving a cooling fluid from a cooling circuit 216, the generally crescent-shaped chamber 218 extending from a front portion 222 positioned near one of a pressure side surface 156 and a suction side surface 158 of the airfoil 150, to a rear portion 154 proximate the trailing edge 154 of the airfoil 150 and the other of the pressure side surface 156 and the suction side surface 158 of the airfoil 150 is positioned, wherein the rear portion 224 of the substantially crescent-shaped chamber 218 in fluid communication with the front portion 222 of the substantially crescent-shaped chamber 218.
    Parts list:
    [0048]<Tb> 100 <September> turbomachinery<Tb> 102 <September> compressor section<Tb> 104 <September> turbine section<Tb> 106 <September> wave<Tb> 108 <September> combustor assembly<Tb> 110 <September> combustion chamber<Tb> 112 <September> Wheels<tb> 114 <SEP> First stage compressor impeller<tb> 116 <SEP> First stage compressor blades<Tb> 118 <September> blade section<Tb> 120 <September> turbine wheels<tb> 122 <SEP> First stage turbine impeller<Tb> 124 <September> turbine blades<Tb> 130 <September> flow path<Tb> 150 <September> blade<Tb> 152 <September> leading edge<Tb> 154 <September> trailing edge<Tb> 156 <September> pressure side surface<Tb> 158 <September> suction surface<tb> 200 <SEP> stationary blades<Tb> 204 <September> end wall<Tb> 212 <September> turbine shroud<Tb> 216 <September> cooling circuit<tb> 218 <SEP> crescent-shaped chamber<Tb> 220 <September> inlets<tb> 222 <SEP> front area<tb> 224 <SEP> rear area<Tb> 226 <September> outlets<tb> 230 <SEP> thermally conductive devices<Tb> 232 <September> containment<Tb> 234 <September> access areas<Tb> 236 <September> transition area<Tb> 238 <September> Lead
    权利要求:
    Claims (10)
    [1]
    A cooling structure for a stationary blade (200) comprising:an end wall (204) connected to a radial end of an airfoil (150) relative to a rotor axis of a turbomachine (100), the airfoil (150) having a pressure side surface (156), a suction side surface (158), a leading edge (152 ) and a trailing edge (154); anda substantially crescent-shaped chamber (218) positioned within the end wall (204) and radially offset from the trailing edge (154) of the airfoil (150), the substantially crescent-shaped chamber (218) containing a cooling fluid from a cooling circuit (216); receives, wherein the substantially crescent-shaped chamber (218) from a front portion (222) which is positioned in the vicinity of one of the pressure side surface (156) and the suction side surface (158) of the airfoil (150) to a rear Portion (224) positioned near the trailing edge (154) of the airfoil (150) and the other of the pressure side surface (156) and the suction side surface (158) of the airfoil (150);wherein the cooling fluid in the front portion (222) is in thermal communication with a portion of the end wall (204) proximate one of the pressure side surface (156) and the suction side surface (158) of the airfoil (150), the cooling fluid in the rear one Region (224) is in thermal communication with a portion of the end wall (204) proximate the trailing edge (154) of the airfoil (150), and wherein the rearward region (224) of the substantially crescent-shaped chamber (218) communicates with the forward region (222 ) of the substantially crescent-shaped chamber (218) is in fluid communication.
    [2]
    The cooling structure of claim 1, wherein the rearward portion (224) of the substantially crescent-shaped chamber (218) extends substantially perpendicularly from one end of the forward portion (222) of the substantially crescent-shaped chamber (218).
    [3]
    The cooling structure of claim 1 or 2, further comprising a plurality of heat conducting means (230) extending through at least one of the front portion (222) and the rear portion (224) of the substantially crescent shaped chamber (218).
    [4]
    The cooling structure of claim 3, wherein the substantially crescent-shaped chamber (218) further includes a perimeter wall (232) and further having a plurality of access areas (234) positioned substantially along the perimeter wall (232) of the substantially crescent-shaped chamber (218) with each of the plurality of access areas (234) being free of heat conducting means (230) therein.
    [5]
    The cooling structure of claim 3 or 4, wherein the substantially crescent-shaped chamber (218) further includes a transition region (236) substantially aligned radially with the trailing edge (154) of the airfoil (150) and between the forward region (222). and the rear portion (224) of the substantially crescent-shaped chamber (218), wherein the transition region (236) is free of heat conducting means (230) therein.
    [6]
    6. The cooling structure of claim 1, wherein an axial length component of the forward portion of the substantially crescent-shaped chamber is at least about half of an axial length component of one of the pressure side surface and the suction side surface of the airfoil (150).
    [7]
    A cooling structure according to any one of the preceding claims, wherein the substantially crescent-shaped chamber (218) further includes a transition region (236) substantially radially aligned with the trailing edge (154) of the airfoil (150) and between the forward region (222 ) and the rear portion (224) of the substantially crescent-shaped chamber (218), and further comprising a projection (238) extending from a radial surface of the transition portion (236), the projection being adapted to engage the projection Directing cooling fluid from the forward region (222) into the rearward region (224) of the substantially sickle shaped chamber (218); andwherein a width of the transition region (236) between the front region (222) and the rear region (224) preferably corresponds approximately to an axial width of the trailing edge (154) of the airfoil (150).
    [8]
    The cooling structure according to any one of the preceding claims, wherein the substantially crescent-shaped chamber (218) comprises one of at least two substantially crescent-shaped chambers (218) positioned within the end wall (204), and wherein the airfoil (150) is one of a pair of airfoils (150) projecting substantially radially from the end wall (204).
    [9]
    9. Stationary bucket (200), comprising:an airfoil (150) including a pressure side surface (156), a suction side surface (158), a leading edge (152) and a trailing edge (154), the airfoil (150) further including a cooling circuit (216) therein;an end wall (204) connected to a radial end of the airfoil (150) relative to a rotor axis of a turbomachine (100); anda substantially crescent-shaped chamber (218) positioned within the end wall (204) and radially offset from the trailing edge (154) of the airfoil (150), the substantially crescent-shaped chamber (218) containing a cooling fluid from the cooling circuit (216); wherein the substantially crescent shaped chamber (218) extends from a forward portion (222) positioned proximate one of the pressure side surface (156) and the suction side surface (158) of the airfoil (150) to a rearward one Portion (224) positioned near the trailing edge (154) of the airfoil (150) and the other of the pressure side surface (156) and the suction side surface (158) of the airfoil (150);wherein the cooling fluid in the front portion (222) is in thermal communication with a portion of the end wall (204) proximate one of the pressure side surface (156) and the suction side surface (158) of the airfoil (150), the cooling fluid in the rear one Region (224) is in thermal communication with a portion of the end wall (204) proximate the trailing edge (154) of the airfoil (150), and wherein the rearward region (224) of the substantially crescent-shaped chamber (218) communicates with the forward region (222 ) of the substantially crescent-shaped chamber (218) is in fluid communication.
    [10]
    A turbine nozzle pair having:a first airfoil (150) having a first cooling circuit (216) therein;an end wall (204) connected to a radial end of the first airfoil (150) relative to a rotor axis of a turbomachine (100);a second airfoil (150) having a second cooling circuit (216) therein, wherein the second airfoil (150) is aligned substantially parallel to the first airfoil, the endwall (204) having a radial end of the airfoil relative to the rotor axis of the turbomachine (100) and wherein each of the first airfoil and the second airfoil further includes a pressure side surface (156), a suction side surface (158), a leading edge (152), and a trailing edge (154);a first substantially crescent-shaped chamber (218) positioned within the end wall (204) and radially offset from the trailing edge (154) of the first airfoil (150), the first substantially crescent-shaped chamber (218) defining a first cooling fluid from the first receives first cooling circuit (216),wherein the first substantially crescent-shaped chamber (218) extends from a front portion (222) positioned near the suction side surface (158) of the first airfoil (150) to a rear portion (224) located in the first Positioned near the trailing edge (154) of the first airfoil (150) and the other of the pressure side surface (156) and the suction side surface (158) of the first airfoil (150), wherein the first cooling fluid in the front region (222) of the first Substantially crescent-shaped chamber (218) is in thermal communication with a portion of the end wall (204) near one of the pressure side surface and the suction side surface of the first airfoil, the first cooling fluid in the rear portion (224) of the first substantially crescent-shaped chamber (218). 218) is in thermal communication with a portion of the end wall (204) proximate the trailing edge (154) of the first airfoil and wherein the rearward portion (224 ) of the first substantially crescent-shaped chamber (218) is in fluid communication with the front portion (222) of the first substantially crescent-shaped chamber (218); anda second substantially crescent-shaped chamber (218) positioned within the end wall (204) and radially offset from the trailing edge (154) of the second airfoil (150), the second substantially crescent-shaped chamber (218) defining a second cooling fluid from the second airfoil second cooling circuit (216) receives,wherein the second substantially crescent-shaped chamber (218) extends from a front portion (222) positioned near the suction side surface (158) of the second airfoil (150) to a rear portion (224) located in the first Near the trailing edge (154) of the second airfoil (150) and the other of the pressure side surface (156) and the suction side surface (158) of the second airfoil is positioned, wherein the second cooling fluid in the front region (222) of the second substantially crescent-shaped chamber (218) is in thermal communication with a portion of the end wall (204) proximate one of the pressure side surface (156) and the suction side surface (158) of the second airfoil (150), the second cooling fluid being in the rearward region (224) of FIG second substantially crescent shaped chamber (218) is in thermal communication with a portion of the end wall (204) proximate the trailing edge (154) of the second airfoil (150) and where at the rear portion (224) of the second substantially crescent-shaped chamber (218) is in fluid communication with the front portion (222) of the second substantially crescent-shaped chamber (218).
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    同族专利:
    公开号 | 公开日
    DE102016112281A1|2017-01-19|
    CN106351699A|2017-01-25|
    JP6835493B2|2021-02-24|
    JP2017025906A|2017-02-02|
    CN106351699B|2020-12-01|
    US9909436B2|2018-03-06|
    US20170016348A1|2017-01-19|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2019-05-31| NV| New agent|Representative=s name: FREIGUTPARTNERS IP LAW FIRM DR. ROLF DITTMANN, CH |
    2019-11-15| AZW| Rejection (application)|
    优先权:
    申请号 | 申请日 | 专利标题
    US14/801,187|US9909436B2|2015-07-16|2015-07-16|Cooling structure for stationary blade|
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