• 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 present invention relates to an airfoil (202) of a gas turbine, the gas turbine having a first sidewall (204), a second sidewall (206), and the airfoil (202) positioned between the first sidewall and the second sidewall. Arranged in the airfoil near a high temperature region is at least one channel configured to direct a cooling fluid to a trailing edge (212) of the airfoil (202), the high temperature region proximate a junction between the first sidewall (204 ) and the trailing edge (212) of the airfoil (202). The airfoil (202) further includes a diffuser (220) in flow communication with the channel, wherein the diffuser (220) is configured to direct the cooling fluid to form a film on a surface of the first sidewall (204), whereby the first sidewall (204) is cooled.
    公开号:CH703886B1
    申请号:CH01592/11
    申请日:2011-09-27
    公开日:2016-07-29
    发明作者:Javier Maldonado Jaime；Michael Itzel Gary
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    Background to the invention
    The subject matter disclosed herein relates to turbines. In particular, the article relates to an airfoil to be positioned in a turbine.
    In a gas turbine, a combustor converts chemical energy of a fuel or an air-fuel mixture into heat energy. The heat energy is conveyed by a fluid, often air from a compressor, to a turbine in which the heat energy is converted into mechanical energy. Several factors influence the efficiency of the conversion of thermal energy into mechanical energy. Factors may include blade pass frequencies, fueling variations, fuel type and reactivity, combustor head volume, fuel nozzle design, air-fuel profiles, flame shape, air-fuel mixing, flame holding, combustion temperature, turbine component design, dilution to mitigate hot gas path temperature, and exhaust gas temperature. For example, For example, high combustion temperatures at selected locations, such as the combustor and turbine nozzle areas, may provide improved combustion efficiency and power generation. In some cases, high temperatures in certain combustor and turbine areas can shorten the life and increase wear and tear on certain components.
    The object underlying the present invention is accordingly to reduce temperatures on selected components of the turbine in order to reduce wear and to increase the service life of turbine components. This object is solved by the subject matter of the independent patent claims. Advantageous developments of the present invention are the subject of the dependent claims.
    Brief description of the invention
    According to the invention, an airfoil which is to be arranged between a first side wall and a second side wall of a gas turbine, a channel through the airfoil, wherein the channel is configured to receive a cooling fluid. The airfoil further includes a diffuser in flow communication with the channel, the diffuser being configured to direct the cooling fluid out of the airfoil to form a film on a surface of the first sidewall, thereby cooling the first sidewall.
    The present invention also relates to a method for cooling a first side wall of a gas turbine using the inventive blade. The method includes directing a cooling fluid to at least one channel in the trailing edge, wherein the cooling fluid is a compressed gas from a compressor, directing the cooling fluid from the at least one channel to a diffuser adjacent the junction between the trailing edge and the first sidewall, and flowing of the cooling fluid from the diffuser to form a film on a surface of the first sidewall, thereby cooling the first sidewall.
    These and other features of the invention will become more apparent from the following description taken in conjunction with the drawings.
    Short description of the drawing
    Features of the invention will become apparent from the following detailed description taken in conjunction with the accompanying drawings, in which:<Tb> FIG. 1 is a schematic drawing of an embodiment of a gas turbine including a combustor, a fuel nozzle, a compressor, and a turbine;<Tb> FIG. FIG. 2 is a perspective view of one embodiment of a turbine nozzle section; FIG.<Tb> FIG. 3 <SEP> is a detailed schematic drawing of one embodiment of a portion of a turbine bucket blade;<Tb> FIG. 4 is a detailed perspective view of one embodiment of a portion of a turbine bucket blade; and<Tb> FIG. 5 is a detailed perspective view of another embodiment of a portion of a turbine bucket blade.
    The detailed description explains embodiments of the invention together with advantages and features by way of example with reference to the drawings.
    Detailed description of the invention
    1 shows a schematic representation of one embodiment of a gas turbine system 100. The system 100 includes a compressor 102, a combustor 104, a turbine 106, a shaft 108, and a fuel nozzle 110. The system 100 may include a plurality of compressors 102, combustors 104 , Turbines 106, shafts 108 and fuel nozzles 110. As shown, the compressor 102 and the turbine 106 are coupled together via the shaft 108. The shaft 108 may be formed by a single shaft or multiple shaft segments coupled together to form the shaft 108.
    The combustor 104 uses liquid and / or gaseous fuel, such as natural gas or a hydrogen-rich syngas, to operate the turbine engine. For example, the fuel nozzles 110 are in fluid communication with a fuel supply and compressed air from the compressor 102. The fuel nozzles 110 create an air-fuel mixture and exhaust the air-fuel mixture into the combustion chamber 104, thereby enabling combustion that generates a hot pressurized exhaust. The combustor 104 directs the hot, pressurized exhaust gas through a transition piece into a turbine nozzle (or "stage 1" nozzle), causing rotation of the turbine 106 as the gas exits the nozzle or vane and onto the nozzle Turbine blade or blade is directed. The rotation of the turbine 106 causes the shaft 108 to circulate, thereby compressing the air as it flows into the compressor 102. In one embodiment, airfoils (also vanes or blades) are disposed in various portions of the turbine, such as compressor 102 or turbine 106, where the gas flow over the airfoils causes wear and thermal fatigue from compressor or compressor due to uneven temperatures Caused turbine components. Controlling the temperature of parts of the turbine bucket blade and adjacent sidewalls can reduce wear and allow a higher combustion temperature in the combustion chamber, thereby improving performance. Cooling of areas near the airfoils and turbine sidewalls will now be described in detail with reference to FIGS. 2-5. Although the following description is primarily directed to gas turbines, the described concepts are not limited to gas turbines.
    FIG. 2 shows a perspective view of one embodiment of a turbine nozzle section 200. The nozzle 200 includes an airfoil 202 positioned between an outer sidewall 204 and an inner sidewall 206. The turbine nozzle 200 receives a hot gas flow 208 from a combustion chamber, the flow causing rotation. The hot gas stream 208 is compressed as it flows past the leading edge 210 and the trailing edge 212 of the airfoil 202. Trailing edge 212 is coupled to outer sidewall 204 and inner sidewall 206 at junctions 214 and 216, respectively. As the hot gas 208 flows over the airfoil 202, cooling channels 219 introduce a cooling fluid 209 into the hot gas, thereby cooling selected portions of the nozzle 200, such as the trailing edge 212. In particular, rows of cooling channels 219 are disposed in the airfoil 202, with the cooling fluid 209 being used to cool the airfoil 202 and sidewalls 204 and 206.
    As shown, the airfoil 202 includes channels 219 disposed at the trailing edge 212. A diffuser 220 is coupled to at least one channel 219 near the junction 214 between the trailing edge 212 and the outer sidewall 214. Similarly, a diffuser 222 is coupled to at least one channel 219 near the junction 216 between the trailing edge 212 and the inner sidewall 216. The diffusers 220 and 222 may be of any suitable configuration and shape to cause the cooling fluid flow to cool an area proximate the junctions 214 and 216. In one embodiment, at least one of the diffusers 220 and 222 is elliptically shaped, as explained below with respect to FIG. 4. In another embodiment, at least one of the diffusers 220 and 222 is triangular in shape, as explained below with reference to FIG. 5. In addition, the geometry of the diffusers 220 and 222 may be described as a contoured opening that promotes formation of a film of cooling fluid on the sidewall 204, 206. As illustrated in FIG. 2, the diffusers 220 and 222 are configured to control a temperature of the surfaces 224 and 226 of the sidewalls 204 and 206, respectively. Additionally, the nozzle 200 may further utilize cooling fluid flow along the sidewall backs 228 and 230 to control a temperature of the sidewalls 204 and 206, respectively.
    Still referring to the embodiment of FIG. 2, the cooling fluid flows from the channels 219 in the airfoil 202, with the channels 219 adjacent the junctions 214 and 216 directing the cooling fluid through the diffusers 220 and 222, respectively. The cooling fluid cools turbine regions or zones of the hot gas path, as well as components of the nozzle 200, such as the airfoil 202 and the sidewalls 204 and 206. The diffusers 220 and 222 are configured to produce a cooling fluid film on the sidewall surfaces 224 and 226, the film the side walls 204 and 206 cools. In addition, the channels 219 of the diffusers 220 and 222 provide convection and conduction cooling at the trailing edge 212. Further, the cooling fluid film isolates the sidewalls 204 and 206 from high temperatures that form in regions proximate the junctions 214 and 216 while the cooling fluid film is cooling Hot gas flows past the blade 202. In embodiments, the cooling fluid is any suitable fluid that cools the nozzle components and selected portions of the gas flow, such as high temperature and high pressure regions within the nozzle. For example, the cooling fluid is a compressed air supply from the compressor, wherein the compressed air is discharged from the supplied air to the combustion chamber. Thus, the cooling fluid is a supply of compressed air that flows around the combustion chamber and is used to cool the turbine nozzle components. The diffusers 220 and 222 located near junctions 214 and 216, respectively, reduce the amount of compressed air used for cooling by improving the cooling of the turbine components and areas in the vicinity of the components. As a result, an increased amount of compressed air is directed to the combustor for conversion to mechanical power output to improve the overall performance and overall efficiency of the turbine engine, thereby extending the life of turbine nozzle parts by reducing oxidation and thermal fatigue. Further, the disclosed means of turbine nozzle 200 and cooling components 219, 220, 222 permit lower temperatures and a more uniform temperature distribution on sidewall 204, 206 and trailing edge 212. In aspects, turbine parts, including the airfoils and side walls, are made of stainless steel or alloy formed, wherein the parts can experience thermal fatigue, if they are not properly cooled during a machine operation. It should be noted that the apparatus may be used in a turbomachine for cooling turbine nozzles, as illustrated in FIGS. 2-5, as well as blades, compressor vanes, or any other airfoils or vanes within a turbomachine.
    3 shows a detailed schematic representation of one embodiment of a portion of a turbine nozzle 300. The turbine nozzle 300 includes a diffuser 302 proximate a joint 304 between a blade air trailing edge 306 and a sidewall 308. A cooling fluid 312 is exhausted from a channel 310 passing the diffuser 302 toward a high temperature region 316, as illustrated by a flow arrow 314. In one embodiment, the high temperature region 316 refers to the turbine components, such as portions of the sidewall 308, as well as an area near the components exposed to elevated temperature and pressure relative to other components in the same region of the turbine. The cooling fluid cools high temperature region 316 and joint 304, and trailing edge 306 and sidewall 308. The hot gas flow from the combustion chamber causes the generation of high temperature and high pressure regions in the nozzle 300 proximate the trailing edge 306 and side wall 308, respectively Diffuser 302 and channel 310 proximate junction 304 enhances cooling of a high temperature region in nozzle 300. The cooling fluid flows through diffuser 302 as illustrated by arrow 314, wherein the flow forms a cooling fluid film on surface 318 of sidewall 308 , The surface 318 may include a thermal barrier coating 320. The thermal barrier coating 320 includes any suitable thermal barrier materials. In a non-limiting example, the thermal barrier coating 320 comprises a metal substrate, a metallic adhesive layer, and a ceramic capping layer. The thermal barrier coating 320 protects turbine components, such as the sidewall 308, from prolonged exposure to heat by using thermally insulating materials that permit a substantial temperature differential between the metallic alloys of the components and the coating surface. Accordingly, the thermal barrier coating 320 allows for higher operating temperatures while limiting the thermal loading of turbine components, such as the sidewall 308. In the illustrated embodiment, the diffuser 302 and the channel 310 are disposed at a location that defines a shoulder 322 that is similar in thickness to the thermal barrier coating 320. When the thermal barrier coating 322 is applied to the sidewall 308, the ledge 322 is filled, thereby providing a smooth transition to the cooling flow 314 as it exits the diffuser 302. This device eliminates additional manufacturing steps to provide the improved joint 304 while allowing the cooling flow 314 to create a cooling fluid film on a surface 318 of the sidewall 308.
    FIG. 4 shows a detailed perspective view of one embodiment of a portion of a turbine nozzle 400. The nozzle 400 includes an elliptical diffuser 402 positioned at or near a junction 404 between the trailing edge 406 and the sidewall 408. The elliptical diffuser 402 is connected to a cooling fluid channel, wherein the cooling fluid flows out of the elliptical diffuser 402 to control a temperature of nozzle parts near the joint 404 and the near high temperature region. The elliptical diffuser 402 may be configured to form a film on a surface 410 of the sidewall 408, wherein the formation of the film cools the surface 410. The cooling fluid channel of the elliptical diffuser 402 also cools the trailing edge 406 by convection and conduction. As shown, the blade trailing edge 406 includes a plurality of channels 412 for cooling the airfoil. In one embodiment, a cooling fluid supply directs pressurized air or any other suitable cooling fluid to multiple passages or channels on the airfoil and back of the sidewall 408, the elliptical diffuser 402 improving cooling of the sidewall 408, trailing edge 406, and joint 404, thereby increasing service life the nozzle components, such as the airfoil and the side wall 408, is extended.
    FIG. 5 shows a detailed perspective view of another embodiment of a portion of a turbine nozzle 500. The nozzle 500 includes a triangular diffuser 502 positioned at a junction 504 between the trailing edge 506 and the sidewall 508. The triangular diffuser 502 is coupled to at least one cooling fluid channel, wherein the cooling fluid flow from the diffuser 502 affects a temperature of nozzle portions near the joint 504 and the nearby high temperature region 512. The airfoil trailing edge 506 includes a plurality of channels 510 to cool the airfoil. It should be noted that the shape of the opening of the diffuser 502 may be any suitable shape for cooling selected portions of the turbine. The shape of the diffuser 502 may be selected based on application specific parameters, manufacturing constraints, and / or costs. In one embodiment, the channels 510 are drilled in the airfoil, and the diffuser 502 is created by electrochemical-mechanical milling or grinding the aperture to the selected shape. In another embodiment, the channels 510 and the diffuser 502 are cast in the selected shapes.
    While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments.
    LIST OF REFERENCE NUMBERS
    [0018]<Tb> FIG. 1 <September><Tb> 100 <September> Turbine System<Tb> 102 <September> compressor<Tb> 104 <September> combustion chamber<Tb> 106 <September> Turbine<Tb> 108 <September> wave<Tb> 110 <September> nozzle<Tb> 112 <September> fuel supply<Tb> <September><Tb> FIG. 2 <September><tb> 200 <SEP> section of a turbine nozzle<Tb> 202 <September> blade<tb> 204 <SEP> outer sidewall<tb> 206 <SEP> inner sidewall<Tb> 208 <September> hot gas flow<Tb> 209 <September> cooling fluid<Tb> 210 <September> leading edge<Tb> 212 <September> trailing edge<tb> 214 <SEP> Junction between trailing edge and sidewall<tb> 216 <SEP> Junction between trailing edge and sidewall<Tb> 219 <September> cooling channels<Tb> 220 <September> diffuser<Tb> 222 <September> diffuser<tb> 224 <SEP> Side wall surface<tb> 226 <SEP> Side wall surface<tb> 228 <SEP> Rear side panel<tb> 230 <SEP> Rear of the side panel<Tb> <September><Tb> FIG. 3 <September><tb> 300 <SEP> section of a turbine nozzle<Tb> 302 <September> diffuser<Tb> 304 <September> junction<Tb> 306 <September> trailing edge<Tb> 308 <September> sidewall<Tb> 310 <September> Channel<Tb> 312 <September> cooling fluid supply<Tb> 314 <September> cooling fluid flow<Tb> 316 <September> high pressure area<Tb> 318 <September> surface<Tb> 320 <September> thermal barrier coating<Tb> <September><Tb> FIG. 4 <September><tb> 400 <SEP> section of a turbine nozzle<tb> 402 <SEP> elliptical diffuser<Tb> 404 <September> junction<Tb> 406 <September> trailing edge<Tb> 408 <September> sidewall<tb> 410 <SEP> Sidewall surface<tb> 412 <SEP> Channels in the trailing edge<Tb> <September><Tb> FIG. 5 <September><tb> 500 <SEP> section of a turbine nozzle<tb> 502 <SEP> triangular diffuser<Tb> 504 <September> junction<Tb> 506 <September> trailing edge<Tb> 508 <September> sidewall<tb> 510 <SEP> Channels in the trailing edge
    权利要求:
    Claims (6)
    [1]
    An airfoil (202) to be disposed between a first (204) and a second sidewall (206) of a gas turbine, the airfoil (202) comprising:a leading edge (210) of the airfoil (202),a trailing edge (212) of the airfoil (202), the trailing edge (212) having at one end a first joint (214) adapted to couple the trailing edge (212) to the first sidewall (204),at least one channel (219), wherein the channel (219) is configured to direct a cooling fluid (209, 312) to the trailing edge (212), andat least one diffuser (220) in flow communication with the channel (219), the diffuser (220) configured to direct the cooling fluid (209, 312) to cool a surface (224, 318) of the first sidewall (204), and shaped to cause a compressed gas forming the cooling fluid (312) to operate to form a film on the surface (224) of the first sidewall (204) to cool the surface (224).
    [2]
    The airfoil (202) of claim 1 having a plurality of channels (219) located at the trailing edge (212), the cooling fluid (312) flowing through the plurality of channels (219, 302) to the trailing edge (212 ) to cool.
    [3]
    The airfoil (202) of claim 1, wherein the diffuser (220) is configured to cool the airfoil trailing edge (212).
    [4]
    The airfoil (202) of claim 1, wherein the diffuser (220) has a triangular (502) or elliptical (402) opening contour.
    [5]
    A method of cooling a first side wall (204, 206) of a gas turbine using an airfoil according to any one of claims 1 to 4, wherein:a cooling fluid (312) is directed to the at least one channel (219, 310) in the trailing edge (212, 306), the cooling fluid (312) being a compressed gas from a compressor (102);the cooling fluid (312) from the at least one channel (219, 310) to the diffuser (220, 222, 302) adjacent the junction (214, 216, 304) between the trailing edge (212, 306) and the first sidewall (204 , 206, 308); andthe cooling fluid (314) flows out of the diffuser (302) to form a film on a surface (224, 226, 318) of the first sidewall (204, 206, 308), whereby the first sidewall (204, 206, 308) is cooled.
    [6]
    The method of claim 5, wherein the cooling fluid (312) is directed to a plurality of channels (219, 310) adjacent the trailing edge (212, 306) and flows through the plurality of channels (218, 310) to the trailing edge (212, 212). 306) to cool.
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2021-04-30| PL| Patent ceased|
    优先权:
    申请号 | 申请日 | 专利标题
    US12/893,506|US8632297B2|2010-09-29|2010-09-29|Turbine airfoil and method for cooling a turbine airfoil|
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