• 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:
    A gas turbine vane section (110) of a gas turbine has an inner end wall (114) with a leading edge (115). A serpentine channel (116) is disposed substantially within the leading edge (115). The serpentine channel (116) has an inlet and an outlet. Air may be taken up at the inlet (118) and discharged while cooling the leading edge (117) at the outlet (119).
    公开号:CH706962B1
    申请号:CH01511/13
    申请日:2013-09-04
    公开日:2017-12-15
    发明作者:Donald Porter Christopher；Jackson Dillard Daniel
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    description
    Technical Field of the Invention The present disclosure relates to a gas turbine vane section and a method of cooling such.
    Background of the Invention Gas turbines, which may also be referred to as combustion turbines, are internal combustion engines that accelerate gases, thereby driving the gases into a combustion chamber in which heat is supplied to them to increase the volume of the gases. The expanding gases are then directed towards a turbine to extract the energy generated by the expanded gases. Gas turbines have many practical applications, including use as jet engines and industrial power plant systems.
    Among the components of a gas turbine, there may be one or more vanes that can direct and accelerate a flow of gases toward the turbine blades (which may also be referred to as "blades") to move the blades about an axis in the center of the turbine Gas turbine to turn to induce. Such vanes may be stationary and in the form of an airfoil extending radially between an outer end wall and an inner end wall of a gas turbine. Hot gases may flow through a path formed by an outer and an inner end wall and the airfoil walls. A leading edge of each end wall may form a protrusion that extends upstream of each such airfoil. As will be appreciated, the leading edge and end walls may reach very high temperatures, and overheating may affect turbine performance. Therefore, measures are often taken to cool these sections of a gas turbine. While the central portion of the end wall may be cooled by impingement cooling, the leading edge portion of a blade airfoil end may structurally overlap portions of a turbine blade (e.g., a turbine blade angel wing seal) to prevent the gases from flowing through the space below the leading edge. This cantilever configuration of the leading edge of a vane end wall may be particularly difficult to cool. This portion may also be subject to particularly high temperatures and stresses, as the blade portions immediately following the leading edge portion of a vane blade may occasionally touch the leading edge. A baffle construction is not well suited for cooling this section, as it is typically not robust enough to survive contact between the blade and the vane blade end wall.
    BRIEF DESCRIPTION OF THE INVENTION According to a first aspect, a gas turbine vane section has an outer end wall and an inner end wall, wherein on the inner end wall, a leading edge provided to the upstream side arrangement is formed. At least one serpentine channel is set up substantially within the leading edge. The serpentine channel has an inlet for receiving air and an outlet for discharging air to cool the leading edge.
    Particularly advantageous and preferred embodiments of the gas turbine vane section according to the first aspect of the present invention have one or more of the following:
    A first portion of the serpentine channel is preferably substantially parallel to a leading edge of the leading edge.
    A second portion of the serpentine channel may be substantially parallel to the leading edge of the leading edge and between the first portion and an airfoil portion of the gas turbine vane portion.
    The at least one serpentine channel includes at least two serpentine channels which are not interconnected.
    The at least one serpentine channel includes at least two serpentine channels of substantially the same size and shape.
    The at least two serpentine channels include two serpentine channels arranged in a substantially mirror-image orientation.
    The at least one serpentine channel is oriented substantially perpendicular to an airfoil of the gas turbine vane section.
    [0005] According to a second aspect, there is provided a method of cooling a gas turbine vane section, wherein air is received at an inlet of a serpentine channel substantially disposed within a leading edge of an inner end wall of the gas turbine vane section, the air is passed through the serpentine channel, and the air is discharged at an outlet of the serpentine channel.
    Particularly advantageous and preferred embodiments of the method according to the second aspect of the present invention have one or more of the following:
    Dispensing the air at an outlet of the serpentine channel comprises discharging the air into a wheel space of a gas turbine.
    Receiving air at an inlet of the serpentine channel includes receiving air at an impingement cooling portion of a gas turbine.
    A first portion of the serpentine channel is preferably substantially parallel to a leading edge of the leading edge.
    A second portion of the serpentine channel may be substantially parallel to the leading edge of the leading edge and between the first portion and an airfoil portion of the gas turbine vane portion.
    The method may further include receiving a second air at a second inlet of a second serpentine channel that is configured substantially within the leading edge of the inner end wall of the gas turbine nozzle portion, passing the second air through the second serpentine channel, and outputting the second air at a second outlet of the second serpentine canal.
    The serpentine canal and the second serpentine canal are preferably not interconnected.
    The serpentine canal and the second serpentine canal may be substantially the same size and shape. The serpentine canal and the second serpentine canal may be arranged substantially mirror-inverted. The serpentine channel may be configured substantially perpendicular to an airfoil of the gas turbine vane section.
    The foregoing summary, as well as the following detailed description, will be better understood when read in conjunction with the drawings. For the purpose of illustrating the claimed subject matter, there are shown in the drawings examples which illustrate various embodiments; however, the invention is not limited to the specific systems and methods disclosed.
    BRIEF DESCRIPTION OF THE DRAWINGS These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings, in which:
    1 shows a side view of a cross-section of a non-limiting exemplary vane and turbine blade portion of a gas turbine engine.
    FIG. 2 is a top plan view of a cross-section of a non-limiting exemplary vane section of a gas turbine engine. FIG.
    3 shows a side view of three cross-sections of a non-limiting example gas turbine nozzle leading edge.
    4 shows a top view of a cross-section of a non-limiting example gas turbine nozzle leading edge.
    5 shows a top view of a cross section of another non-limiting example gas turbine nozzle leading edge.
    6 shows a top view of a cross section of another non-limiting example gas turbine nozzle leading edge.
    7 shows a top view of a cross section of another non-limiting example gas turbine nozzle leading edge.
    8 shows a top view of a cross section of another non-limiting example gas turbine nozzle leading edge.
    9 shows a top view of a cross section of another non-limiting example gas turbine nozzle leading edge.
    DETAILED DESCRIPTION OF THE INVENTION FIG. 1 illustrates a non-limiting example portion 100 of a gas turbine having a vane portion and a blade portion shown in a cross-sectional side view. It should be noted that FIG. 1 shows a simplified illustration of a portion of a gas turbine for illustrative purposes and omits many components and systems that may be present in accordance with any of the embodiments discussed herein. As one skilled in the art will appreciate, the embodiments disclosed herein can be implemented in many different gas turbines of any type and construction. All such embodiments are considered to be within the scope of the present disclosure.
    The section 100 includes a turbine blade 120 containing an airfoil 122. The turbine blade 120 may rotate about the axis 140. The section 100 also includes a vane section 110 that is stationary and has an outer end wall 113 and an inner end wall 114. Between the end walls 113 and 114 and connected to them, the airfoil 112 is arranged. The inner end wall 114 has a leading edge 115 that is upstream of the airfoil 112. An airflow within the section 100 flows in the direction 130, passing over and through the airfoil 120 and then over and through the vane section 110.
    In order to minimize the flow of gases through the space below the leading edge 115, the turbine blade 120 has a portion 124 (eg, an "angel wing seal") configured to seal between the turbine blade 120 and the vane portion 110 to create. While the central portion 117 of the inner end wall 114 may be cooled by an impact when air bypasses the airfoil 112, it may be difficult to cool the leading edge 115 of the vane portion 110. Accordingly, in one embodiment, one or more channels 116 are disposed substantially within the leading edge 115 to receive air at a channel inlet 118 after this air has caused impingement cooling of the central portion 117 of the inner end wall 114 (the air flow direction in FIG Arrows is indicated). The air received via the inlet 118 cools the leading edge 115 by flowing through the channels 116. The channels 116 are configured such that the air is discharged into the wheel space of the gas turbine section 100 through the channel outlet 119 after passing through the channels 116. It should be noted that any configuration, construction, number and location of inlets and outlets of channels, such as channels 116, are considered to be within the scope of the present disclosure.
    FIG. 2 illustrates a top view of a cross section of a vane section 200, including the leading edge 210, according to one embodiment. It should be noted that Figs. 2-9 are not drawn to scale and those portions for which a description exists are shown in more detail. A pair of channels 230 and 240 are arranged at the front edge 210. The airfoil 220 is shown in phantom to provide a baseline for the location of the channels and the orientation of the vane section 200. As seen in this figure, the channels 230 and 240 are within the leading edge 210. It should be noted that although FIG. 2 illustrates a single airfoil for purposes of example, a plurality of airfoils may be present in each vane section in which each of the airfoils present embodiments, and all such embodiments are considered to be within the scope of the present disclosure.
    In one embodiment, the channels 230 and 240 have substantially the same size and shape and are serpentine, with a first portion of the serpentine channel along the leading edge or the outermost portion of the leading edge 210, i. farthest from the airfoil 220, and then a second portion of the serpentine channel is closer to the location of the airfoil 220. In this way, the leading edge portion can be cooled by the channels. The first and second portions may be connected to other portions to form a total of serpentine channels 230 and 240, as shown in FIG. For example, air may enter each of the channels 230 and 240 through inlet passages 231 and 241, respectively (arrows indicating the air flow are illustrated in FIG. 2) through which portions of the passages 230 and 240 pass and then at outlet passages 232 and 242 of FIG Pass channels 230 and 240, respectively. The air received at the inlet passages 231 and 241 may be received by an impingement cooling area of a vane portion or an impact area of an end wall. Alternatively, the air received at the inlets may be received by any other portion of a gas turbine, including an impingement cooling area and a non-impingement cooling area of an airfoil. It should be understood that outlets, inlets, and associated portions may be located at any location on the leading edge 210 or elsewhere in the vane section 200, and multiple outlets and / or inlets per channel will be within the scope of the present disclosure considered. For example, as shown in FIG. 2, an inlet may be disposed toward the airfoil portion of a vane portion (for further examples, see also FIG. 1, inlet 118 and FIG. 3, inlet 311) and configured to provide air after impingement cooling the end wall while an outlet may be disposed toward the wheel space (for further examples see also Fig. 1, outlet 119 and Fig. 3, outlet 331). It should be noted that in some embodiments, the channels 230 and 240 may be interconnected, while in other embodiments, the channels 230 and 240 may not be interconnected.
    FIG. 3 illustrates three cross-sectional side views of a leading edge having a serpentine channel, such as that shown in FIG. 2. A section 310 shows a side view of a cross-section of the inlet portion of the serpentine channel, in which the inlet 311 allows air to enter the channel and to flow through the channel portion 312 to the outermost portion of the leading edge. The section 320 shows a side view of the cross section of the central portion of the serpentine channel in which air flows through the channel portion 312 to the channel portion 321, which is closer to the airfoil portion of the vane portion than the channel portion 312. The section 330 shows a cross-sectional side view of the outlet section of the serpentine channel, in which the outlet 331 discharges the air flowing from the channel section 321, through the outlet 331, to the wheel space of the turbine. As seen in this figure, an outlet is disposed at a side or at a lower portion of the leading edge so that air can be discharged into the wheel space of the turbine, and an inlet at the side of the leading edge in an area near an impingement cooled portion a vane section arranged. In other embodiments, the outlet and / or the inlet may each be located elsewhere, and all such embodiments are considered to be within the scope of the present disclosure.
    4-9 illustrate modified exemplary embodiments of serpentine channels that may be used to cool a leading edge of a gas turbine nozzle vane section. It is to be understood that such channels in these or any other embodiments described herein may have any shape, size, number, orientation, configuration, and all such embodiments are considered to be within the scope of the present disclosure.
    Fig. 4 illustrates an embodiment in which the channels have approximately the same shape as those of Fig. 2, but with different channel sizes. It should be noted that only the leading edge portion of an inner end wall of a gas turbine vane portion is shown in Figs. 4-9 so as to provide a clear explanation of the intended embodiments.
    In Fig. 4, the leading edge 400 has serpentine channels 430 and 440, each having a first portion of the serpentine channel extending along the outermost, i.e., the first, portion of the serpentine channel. farthest from the airfoil, near the leading edge 400 of the leading edge 400, and then provided with a second portion of the serpentine channel closer to the location of the associated airfoil. In this embodiment, the inner diameter, or width, of the channel may vary. For example, the widths 433 and 443 of the first portion of the channels 430 and 440, respectively, may be smaller than the widths 434 and 444 of the second portion of the channels 430 and 440, respectively. Alternatively, in another embodiment, the width of a first portion may be greater than that of a second portion Section of a channel. Any width, diameter, dimensions of any channel and their variations are considered to be within the scope of the present disclosure. Similar to the portions of FIG. 2, the leading edge 400 may be cooled by the channels 430 and 440 as air enters each of the channels 430 and 440 through the inlets 431 and 441, respectively (arrows indicating the airflow are shown in FIG illustrated, while the complete inlet section and outlet section, as shown in Fig. 2 are not shown for clarity), flows through the channels and then exits at the outlets 432 and 442 of the channels 430 and 440, respectively. It should be noted that, as with all embodiments disclosed herein, the outlets and the inlets may be located anywhere on the leading edge or elsewhere in the vane section, and multiple outlets and / or inlets per channel are considered to be circumferential considered in accordance with the present disclosure.
    FIG. 5 illustrates a non-limiting example leading edge 500 having serpentine channels 530 and 540, each having more than two sections, unlike the embodiment illustrated in FIGS. 2 and 4. First portions 535 and 545 of channels 530 and 540, respectively, run along the outermost portion of leading edge 500, i. farthest from one airfoil, near the leading edge 500, while second passages 536 and 546 of the passages 530 and 540, respectively, are centrally located between the other two portions, third portions 537 and 547, respectively, of the passages 530 and 540 at the Location of an associated airfoil closest. In this embodiment, as in all other disclosed herein, the inner diameter or width of the channel may vary. Any number of sections and any configurations of such sections are considered to be within the scope of the present disclosure. By using multiple sections per channel established in the leading edge 500, the leading edge 500 can be cooled more easily and more thoroughly through the channels 530 and 540 as air enters each of the channels 530 and 540 via the inlets 531 and 541, respectively (arrows indicating airflow 5) through which channels pass and then exit at the outlets 532 and 542 of the channels 530 and 540, respectively. It should be noted that, as with all of the embodiments disclosed herein, the outlets and the inlets may be located at any location on a leading edge or elsewhere in a vane section and multiple outlets and / or inlets per channel than within the scope of the present invention Revelation be considered lying.
    FIG. 6 illustrates a non-limiting exemplary leading edge 600 that may include more than two serpentine channels, illustrating that each number of channels is considered to be within the scope of the present disclosure. In the illustrated embodiment, three channels 630, 640, and 650 may all be arranged in a leading edge 600, and each of them may have any number of sections, each of which may be any diameter or width. Any number of sections and any configurations of such sections for each of the channels 630, 640 and 650 are considered to be within the scope of the present disclosure. By using multiple sections in the leading edge 600, the leading edge 600 can be more easily and thoroughly cooled (arrows indicating airflow are illustrated in FIG. 6).
    Instead of multiple channels, in one embodiment in a leading edge, a single and optionally longer channel can be used. Fig. 7 illustrates such an embodiment. The leading edge 700 has a single channel 730, which may have any number of sections, each of which may have any diameter or width. Any number of sections and any configurations of such sections for channel 730 are considered to be within the scope of the present disclosure. By using a single and optionally longer portion in the leading edge 700, the passageway 730 can thoroughly cool the leading edge 700 (arrows indicating airflow are illustrated in FIG. 7).
    Any orientation of the channels relative to each other and to any other part of a nozzle section of a gas turbine is considered to be within the scope of the present disclosure. For example, For example, as shown in FIG. 8, each of the channels 830 and 840 in the leading edge 800 may be oriented in mirror image orientation. In another example shown in FIG. 9, each of the channels 930 and 940 in the leading edge 900 is constructed such that the longer portions of the channels are substantially perpendicular to the front side (ie, the portion farthest from an associated airfoil). of the leading edge, rather than being substantially parallel to the front of a leading edge, as in FIGS. 4, 7 and 8, for example.
    The technical effect of the systems and methods described herein is improved cooling of a leading edge portion of a gas turbine vane by using impingement cooling air. As will be appreciated by those skilled in the art, the use of the disclosed processes and systems can reduce the temperature or required cooling flow on a substantial gas turbine component and therefore improve component performance and life. The fabrication of the disclosed embodiments may also eliminate some post-casting operations, which may result in cost savings. Those skilled in the art will recognize that the disclosed end-wall cooling systems and methods may be combined with other cooling systems and technologies to achieve even greater temperature or cooling flow reduction. All such embodiments are considered to be within the scope of the present disclosure.
    REFERENCE SIGNS LIST 100 gas turbine section 1 110 vane 1 112 airfoil 1 113 outer end wall 1 114 inner end wall 1 115 leading edge 1 116 channels 1 117 end wall section 1 118 channel inlet 1 119 channel outlet 1 120 turbine blade 1 122 airfoil 1 124 turbine blade section 1 140 axis 1 200 vane section 2 210 Front edge 2 230 Channel 2 231 Channel inlet 2 232 Channel outlet 2 240 Channel 2 241 Channel inlet 2 242 Channel outlet 2 310 Leading edge section 3 311 Channel inlet 3 312 Channel section 3 320 Leading edge section 3 321 Channel section 3 330 Leading edge section 3 331 Channel outlet 3 400 Leading edge 4 430 Channel 4 432 Duct outlet 4 433 Duct width 4 434 Duct width 4 440 Duct 4 442 Duct outlet 4 443 Duct width 4 444 Duct width 4 500 Front edge 5 530 Duct 5 531 Duct inlet 5 532 Duct outlet 5 535 Duct passage 5 536 Duct passage 5 537 Duct passage 5 540 Duct 5 541 Duct inlet 5 542 Duct outlet 5 545 canal passage 5 546 canal passage ng 5 547 Channel passage 5 600 Front edge 6 630 Channel 6 640 Channel 6 650 Channel 6 700 Front edge 7 730 Channel 7 800 Front edge 8 830 Channel 8 840 Channel 8 900 Front edge 9
    权利要求:
    Claims (10)
    [1]
    930 channel 9 940 channel 9 claims
    A gas turbine vane section (110) comprising: an outer end wall (113) and an inner end wall (114), wherein a leading edge (115/210/400/500/600/700/800/900 ), which is provided to the upstream side arrangement, and wherein at least one serpentine channel (116/230/240/312/321/430/440/530/540/630/640/650/730/830/840/930/940 ) is set up substantially within the leading edge (115/210/406/500/600/700/800/900), wherein the serpentine channel (116/230/240/312/321/430/440/530/540/630 / 640/650/730/830/840/930/940) has an inlet (118/231/241/311/531) for receiving air and an outlet (119/232/242/331/532) for discharging air ,
    [2]
    The gas turbine vane section (110) of claim 1, wherein a first portion (535/545) of the at least one serpentine channel (530/540) is substantially parallel to a leading edge of the leading edge (500).
    [3]
    The gas turbine vane section (110) of claim 2, wherein a second section (536/546) of the serpentine channel (530/540) is substantially parallel to the leading edge of the leading edge (500) and between the first section (535/545) and a second section Airfoil portion (122) of the gas turbine vane portion (110) extends.
    [4]
    Gas turbine vane section (110) according to any one of the preceding claims, wherein at least two serpentine channels (230/240/430/440/530/540/630/640/650/830/880/930/940) are connected to the at least one serpentine channel (230 / 240/430/440/530/540/630/640/650/830/880/930/940), which are not interconnected.
    [5]
    5. Gas turbine guide vane section (110) according to one of the preceding claims, wherein at least one serpentine channel (230/240/430/440/530/540/630/640/650/930/940) at least two serpentine channels (230/240/430 / 440/530/540/630/640/650/930/940) which are substantially the same size and shape.
    [6]
    A gas turbine vane section (110) according to any one of the preceding claims, wherein said at least one serpentine channel (830/840) includes at least two serpentine channels (830/840) arranged substantially mirror-inverted.
    [7]
    A gas turbine vane section (110) according to any one of the preceding claims, wherein the at least one serpentine channel is oriented substantially perpendicular to an airfoil (122) of the gas turbine vane section (110).
    [8]
    A method of cooling a gas turbine vane section (110) of a gas turbine, comprising: receiving air at an inlet (118/231/241/311/531) of a serpentine channel (116/230/240/312/321/430/440 / 530/540/630/640/650/730/830/840/930/940) substantially within a leading edge (115/210/400/500/600/700/800/900) of an inner end wall (114). the gas turbine vane section (110) is arranged; Passing the air through the serpentine duct (116/230/240/312/321/430/440/530/540/630/640/650/730/830/840/930/940); and discharging the air at an outlet (119/232/242/331/532) of the serpentine channel (116/230/240/312/321/430/440 / 530/540/630/640/650/730/830/840 / 930/940).
    [9]
    A method according to claim 8, wherein air is discharged from the outlet (119/232/242/331/532) into a wheel space of the gas turbine.
    [10]
    10. The method of claim 8 or 9, wherein at the inlet (118/231/241/311/531) air is received from an impingement cooling section of the gas turbine.
    类似技术:
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    同族专利:
    公开号 | 公开日
    US20140072400A1|2014-03-13|
    JP2014051981A|2014-03-20|
    DE102013109150A1|2014-03-13|
    CH706962A2|2014-03-14|
    CH706962A8|2014-08-29|
    US9194237B2|2015-11-24|
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    法律状态:
    2014-08-29| PK| Correction|Free format text: ERFINDER BERICHTIGT. |
    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
    US13/608,269|US9194237B2|2012-09-10|2012-09-10|Serpentine cooling of nozzle endwall|
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