• 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 a turbine blade cooling system (100). The turbine blade cooling system (100) includes a first turbine blade (120) having a first turbine blade platform (150) with a cooling cavity in communication with a pressure side channel and a second turbine blade (130) with a second turbine blade platform (150) having a platform cooling cavity with a suction side channel. The pressure side channel of the first turbine blade platform (150) communicates with the suction side channel of the second turbine blade platform (150).
    公开号:CH703763B1
    申请号:CH01436/11
    申请日:2011-09-01
    公开日:2016-01-15
    发明作者:Bradley Taylor Boyer
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    Technical area
    The present application relates to a turbine blade platform cooling system and to a method of cooling turbine blades.
    Background to the invention
    Known turbine assemblies generally include rows of circumferentially spaced turbine blades. Generally described, each turbine blade has an airfoil extending outwardly from a platform and a dovetailed shaft extending inwardly from the shaft. The dovetail is used to attach the turbine blade to a rotor disk for rotation therewith. Known turbine blades are generally hollow so that an internal cooling cavity may be formed through at least portions of the airfoil, the platform, the shaft and the dovetail.
    There may be temperature differences at the transition between the airfoil and the platform and / or between the shaft and the platform because the airfoil portions of the blades are exposed to higher temperatures than the stem and dovetail portions. Over time, such temperature differences and the associated thermal stresses can exert high compressive thermal stresses on the blade platform. In addition, the increased operating temperatures of the turbine as a whole may cause oxidation, fatigue, breakage, and / or creep flexing, and thereby a shorter useful life of the turbine blade. Specifically, the potential stresses of the entire turbine blade and blade platform in general increase with higher combustion temperatures of the turbines.
    There is therefore a desire for a turbine blade with improved cooling in particular around the suction side of the platform. Such improved turbine blade design would allow the use of higher combustion temperatures and thereby higher overall system efficiency with extended component life.
    Brief description of the invention
    The present invention relates to a turbine blade cooling system. The turbine blade cooling system includes a first turbine blade having a first turbine blade platform with a cooling cavity in communication with a pressure side channel, and a second turbine blade having a second turbine blade platform with a platform cooling cavity having a suction side channel. The pressure side channel of the first turbine blade platform communicates with the suction side channel of the second turbine blade platform.
    The present invention further relates to a method of cooling a turbine blade using a turbine blade cooling system according to the invention. The method comprises the steps of: passing a coolant through the pressure side channel of the first turbine blade platform and into the suction side channel and the platform cooling cavity of the second turbine blade platform; Passing the coolant through the suction side channel of the second turbine blade platform; and directing the coolant through the platform cooling cavity in the second turbine blade platform.
    These and other features and improvements of the present application will become apparent to those skilled in the art upon review of the following detailed description, taken in conjunction with the several drawings and the appended claims.
    Brief description of the drawings
    [0008]<Tb> FIG. 1 <SEP> is a schematic view of the components of a known gas turbine plant.<Tb> FIG. 2 <SEP> is a perspective view of a known turbine blade.<Tb> FIG. 3 <SEP> is a plan view of a pair of turbine blades of the turbine blade platform cooling system as described herein.<Tb> FIG. 4 <SEP> is a side cross-sectional view of the pair of turbine blades of the turbine blade platform cooling system of FIG. 3.<Tb> FIG. 5 is a partial side perspective view of the pair of turbine blades of the turbine blade platform cooling system of FIG. 3 shown separated.
    Detailed description of the invention
    Referring now to the drawings, wherein like reference numbers refer to like elements throughout the several views: FIG. 1 shows a schematic view of the components of a known gas turbine plant 10. The gas turbine plant 10 may include a compressor 15. The compressor 15 compresses an incoming air stream 20. The compressor 15 supplies the compressed air stream 20 to a combustion chamber 25. The combustor 25 mixes the compressed air stream 20 with a compressed fuel stream 30 and ignites the mixture to produce a combustion gas stream 35. While only a single combustor 25 is shown, the gas turbine engine 10 could include any number of combustors 25. The combustion gas stream 35 is in turn fed to a turbine 40. The combustion gas stream 35 drives the turbine 40 to perform mechanical work. The mechanical work done in the turbine 40 drives the compressor 15 and an external load 45, e.g. an electric generator and the like.
    The gas turbine plant 10 may use natural gas, various types of syngas, or other types of fuels. The gas turbine plant 10 may be one of any number of different gas turbines offered by General Electric Company of Schenectady, New York, or others. The gas turbine plant 10 may have a different structure and use other types of components. Other types of gas turbine plants could also be used herein. Several gas turbine plants 10 or other types of turbines and other types of power plants could be shared herein.
    Fig. 2 shows a perspective view of a known turbine blade 50. The turbine blade 50 may be used in the turbine 40 as described above and the like. Any number of the blades 50 may be disposed adjacent each other in a circumferentially spaced row. Each turbine bucket 50 generally includes an airfoil 55 extending from a platform 60. The airfoil 55 may be convex in shape having a suction side 65 and a pressure side 70. Each airfoil 55 may also have a leading edge 75 and a trailing edge 80. Other airfoil designs could be used herein.
    The turbine blade 50 may also include a stem 85 and a dovetail 90 extending inwardly from the platform 60. A number of angel wings 86 may be attached to the stem 85. The dovetail 90 may connect the turbine blade 50 to a disk (not shown) for rotation therewith. The shaft 85 may be substantially hollow with a stem cavity 95 therein. The shaft cavity 95 may be filled with a coolant, such as a coolant. Compressor outlet air communicate. Other types of refrigeration circuits and refrigerants may also be used herein. The coolant may circulate through at least portions of the dovetail 90, the stem 85, the platform 60 and into the airfoil 55. Other configurations could be used herein.
    Figures 3 to 5 show a turbine blade platform cooling system 100 as described herein. The turbine blade platform cooling system 100 may include any number of turbine blades 110, with only a first turbine blade 120 and a second turbine blade 130 shown. As described above, any number of turbine buckets 110 may be circumferentially disposed adjacent each other about a rotor disk (not shown). Each pair of turbine blades 110 may form a gap 140 therebetween. The first turbine blade 120 and the second turbine blade 130 may be substantially identical.
    Each turbine blade 110 has a platform 150 with an airfoil 160 extending outwardly therefrom and a shaft 170 extending inwardly from the platform. The platform 150 has a front side 152, a rear side 154, a suction side 156 and a pressure side 158.
    The turbine blade 110 has a cooling cavity 180 extending therethrough. The cooling cavity 180 may be equipped with a coolant 190, such as a coolant. Compressor outlet air and the like are in communication. The cooling cavity 180 may extend at least partially through the shaft 170 and into the airfoil 160. A portion of the cooling cavity 180 may also extend into the platform 150 such that at least a portion of the coolant 190 may flow through the platform, either through or past the airfoil 160. Specifically, the cooling cavity 180 may extend into the rear portion 154 of the platform 150 about the pressure side 158 thereof. The portion of the cooling cavity 180 terminates around a pressure side channel 200 of the platform 150. Other configurations could be used herein.
    The platform 150 also includes a platform cooling cavity 210. The platform cooling cavity 210 extends from the suction side 156 of the platform 150 to the rear side 154. The platform cooling cavity 210 begins around a suction side channel 220. The suction side channel 220 is aligned with the pressure side channel 200 of the adjacent turbine blade 110 to pass the coolant 190 through these channels. The platform cooling cavity 210 may also include a rear port 230 for discharging the coolant 190 as it passes therethrough. The platform cooling cavity 210 may also include a pin assembly or other types of turbulators therein to create turbulence for increased heat transfer. Other types of internal configurations could be used herein.
    In operation, the coolant 190 flows through the cooling cavity 180 of the first turbine blade 120. At least a portion of the coolant 190 flows through the platform 150 and exits through the pressure side channel 200. The coolant 190 then flows through the gap 140 and into the platform cooling cavity 210 of the second turbine blade 130. Specifically, the coolant 190 flows into the suction side channel 220 of the platform cooling cavity 210 located on the suction side 156 of the platform 150 along the rear end 154 thereof. The coolant 190 may then exit the platform 150 along the rearward channel 230.
    The turbine blade platform cooling system 100 thus permits cooling on the suction side 156 of the platform 150 of the second turbine blade 130 by the coolant 190 from the first turbine blade 120. The pin assembly or other types of turbulators 240 within the platform cooling cavity 210 also allow for increased heat transfer therein , This cooling also allows some lateral flexibility between the cooler shaft side and the hot gas side of the platform 150 to reduce the thermal stresses therein. Surface film openings and the like may also be used herein in conjunction with the platform cooling cavity 210. Various types of seals around the gap 140 may also be used to reduce leakage or inflow therethrough.
    The turbine blade platform cooling system 100 thus provides platform cooling to allow for higher turbine operating temperatures to allow for higher efficiencies and lower customer operating costs with less effect on the durability of the components. Using the coolant 190 from the first blade 120 to cool the second blade 130 further increases the overall efficiency. A transition of the coolant 190 may also be made from the suction side 156 to the pressure side 158 in a similar manner. Any type of cooling system could be used herein from platform to platform in any direction.
    The present invention relates to a turbine blade cooling system 100. The turbine blade cooling system 100 includes a first turbine blade 120 having a first turbine blade platform 150 with a cooling cavity 180 in communication with a pressure side channel 200 and a second turbine blade 130 with a second turbine blade platform 150 having a platform cooling cavity 210 with one Suction side channel 220. The pressure side channel 200 of the first turbine blade platform 150 communicates with the suction side channel 220 of the second turbine blade platform 150.
    LIST OF REFERENCE NUMBERS
    [0021]<Tb> 10 'September> gas turbine plant<Tb> 15 <September> compressor<Tb> 20 <September> airflow<Tb> 25 <September> combustion chamber<Tb> 30 <September> fuel stream<Tb> 35 <September> combustion gas stream<Tb> 40 <September> Turbine<Tb> 45 <September> Last<Tb> 50 <September> turbine blade<Tb> 55 <September> blade<Tb> 60 <September> Platform<Tb> 65 <September> suction<Tb> 70 <September> Print Page<Tb> 75 <September> leading edge<Tb> 80 <September> trailing edge<Tb> 85 <September> End<Tb> 86 <September> angel wings<Tb> 90 <September> Swallowtail<Tb> 95 <September> shank cavity<Tb> 100 <September> turbine blade cooling system<Tb> 110 <September> turbine blade<tb> 120 <SEP> First turbine blade<tb> 130 <SEP> Second turbine blade<Tb> 140 <September> gap<Tb> 150 <September> Platform<Tb> 152 <September> Front<Tb> 154 <September> back<Tb> 156 <September> suction<Tb> 158 <September> Print Page<Tb> 160 <September> blade<Tb> 170 <September> End<Tb> 180 <September> cooling cavity<Tb> 190 <September> coolant<Tb> 200 <September> pressure side channel<Tb> 210 <September> platform cooling cavity<Tb> 220 <September> Saugseitenkanal<Tb> 230 <September> rear channel<Tb> 240 <September> turbulator
    权利要求:
    Claims (14)
    [1]
    A turbine blade cooling system (100) comprising:a first turbine blade (120);wherein the first turbine blade (120) includes a first turbine blade platform (150) and a cooling cavity (180);wherein the cooling cavity (180) communicates with a pressure side channel (200) in the first turbine blade platform (150); anda second turbine blade (130);wherein the second turbine bucket (130) includes a second turbine bucket platform (150) and a platform cooling cavity (210);wherein the platform cooling cavity (210) has a suction side channel (220) in communication with the pressure side channel (200).
    [2]
    The turbine blade cooling system (100) of claim 1, wherein the first turbine blade platform (150) has a pressure side (158) with the pressure side channel (200) disposed therein.
    [3]
    The turbine blade cooling system (100) of claim 1, wherein the first turbine blade platform (150) has a rearward side (154) axially downstream with respect to its installed state in a turbine and the pressure side channel (200) is disposed therein.
    [4]
    The turbine blade cooling system (100) of claim 1, wherein the second turbine blade platform (150) has a suction side (156) and the suction side channel (220) is disposed therein.
    [5]
    The turbine blade cooling system (100) of claim 1, wherein the second turbine blade platform (150) has a suction side (156) and the platform cooling cavity (210) is disposed therein.
    [6]
    The turbine blade cooling system (100) of claim 1, wherein the second turbine blade platform (150) has a rearward side (154) axially downstream with respect to its installed state in a turbine and the platform cooling cavity (210) is disposed therein.
    [7]
    The turbine blade cooling system (100) of claim 1, wherein the second turbine blade platform (150) has an axially downstream rear face (154) with respect to its installed state in a turbine, and the platform cooling cavity (210) has a rearward channel (230) thereon.
    [8]
    The turbine blade cooling system (100) of claim 1, further comprising a gap (140) between the first turbine blade platform (150) and the second turbine blade platform (150) in the installed state of the first and second turbine blade platforms (150) in a turbine.
    [9]
    The turbine blade cooling system (100) of claim 8 including seals around said gap (140) to reduce leakage or inflow through said gap (140).
    [10]
    The turbine blade cooling system (100) of claim 1, wherein the platform cooling cavity (210) includes a number of turbulators (240) therein.
    [11]
    11. A method of cooling first and second turbine blades (120, 130) using a turbine blade cooling system (100) according to any one of claims 1 to 10, comprising:Passing a coolant (190) through the pressure side channel (200) of the first turbine blade platform (150) and into the suction side channel (220) and the platform cooling cavity (210) of the second turbine blade platform (150);Passing the coolant (190) through the suction side channel (220) of the second turbine blade platform (150); andPassing the coolant (190) through the platform cooling cavity (210) in the second turbine bucket platform (150).
    [12]
    12. The method of claim 11, wherein the step of directing the coolant (190) through the platform cooling cavity (210) comprises generating turbulence therein.
    [13]
    13. The method of claim 11, further comprising the step of directing the coolant (190) through the rearward channel (230) out of the platform cooling cavity (210).
    [14]
    14. The method of claim 11, further including the step of directing the coolant (190) through the airfoil (160) of the first turbine blade (120).
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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
    US12/878,075|US9416666B2|2010-09-09|2010-09-09|Turbine blade platform cooling systems|
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