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公开(公告)号：EP3088100B1
公开(公告)日：2020-01-01
申请号：EP16167665.5
申请日：2016-04-29
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公开(公告)号：EP3211179A1
公开(公告)日：2017-08-30
申请号：EP17158004.6
申请日：2017-02-24
IPC分类号： F01D5/18
CPC分类号： F01D5/187 , B22C9/101 , F01D9/041 , F01D11/08 , F01D25/12 , F05D2220/32 , F05D2230/20 , F05D2230/211 , F05D2240/122 , F05D2240/304 , F05D2240/55 , F05D2260/2214 , F05D2260/22141 , Y02T50/676
摘要： An airfoil (201) of a gas turbine engine includes an airfoil body (203) having a leading edge (212) and a trailing edge (214) extending in a radial direction, a trailing edge cavity (230) formed within the airfoil and proximate to the trailing edge of the airfoil, the trailing edge cavity extending from the trailing edge in a forward direction toward the leading edge, at least one set of blocking pedestals (240; 242) located within the trailing edge cavity, a set of circular pedestals (244) located aftward from the at least one blocking set of pedestals, and a set of spear pedestals (246) located aftward from the set of circular pedestals and closest to the trailing edge of the airfoil body. A corresponding core for manufacturing an airfoil is also provided.
摘要翻译： 燃气涡轮发动机的翼型（201）包括具有在径向方向上延伸的前缘（212）和后缘（214）的翼型体（203），在翼型内形成并且紧邻的后缘腔（230） 所述后缘腔从所述后缘朝前缘向前延伸，至少一组位于所述后缘腔内的阻挡基座（240; 242），一组圆形基座 （244），其位于所述至少一个阻塞组底座的后方;以及一组矛形底座（246），其位于所述一组圆形底座的后方并且最靠近所述翼型体的所述尾缘。 还提供了用于制造翼型的相应的芯。
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公开(公告)号：EP3088100A1
公开(公告)日：2016-11-02
申请号：EP16167665.5
申请日：2016-04-29
摘要： A casting core (196) for an airfoil (165) includes a first portion (197) extending in a first direction and corresponding to a first cavity (179, 180) of an airfoil (165). The first portion (197) defines a reference plane along a parting line (204) formed by a casting die (202). The first portion (197) defines a plurality of grooves (212) corresponding to a plurality of trip strips (195) of the airfoil (165). Each of the plurality of grooves (212) defines a respective groove axis (214), and the plurality of grooves (212) are distributed in the first direction along a first side of the reference plane such that one or more of the groove axes (214) are oriented with respect to a pull direction (P) of the casting die (202). A method for fabricating a gas turbine engine component is also disclosed.
摘要翻译： 用于翼型件（165）的铸造芯（196）包括沿第一方向延伸并且对应于翼型件（165）的第一空腔（179,180）的第一部分（197）。 第一部分（197）沿着由铸模（202）形成的分型线（204）限定参考平面。 第一部分（197）限定对应于翼型件（165）的多个跳闸条（195）的多个凹槽（212）。 多个凹槽（212）中的每一个限定相应的凹槽轴线（214），并且多个凹槽（212）沿着参考平面的第一侧沿第一方向分布，使得一个或多个凹槽轴线 214）相对于铸模（202）的拉动方向（P）定向。 还公开了一种用于制造燃气涡轮发动机部件的方法。
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4.发明公开GAS TURBINE ENGINE AIRFOIL COOLING PASSAGE CONFIGURATION 审中-公开KONFIGURATION DESKÜHKKALALSEINER GASTURBINENMOTORSCHAUFEL
公开(公告)号：EP2937511A1
公开(公告)日：2015-10-28
申请号：EP15163567.9
申请日：2015-04-14
IPC分类号： F01D5/18
CPC分类号： F01D5/186 , F01D5/187 , F05D2250/185 , Y02T50/676
摘要： A component (64) for a gas turbine engine includes an exterior surface that provides pressure and suction sides (86,88). A cooling passage (90) in the component (64) includes a serpentine passageway (94) that has first and second passes (118,120) respectively configured to provide fluid flow in opposite directions from one another. The first pass (118) includes first and second portions (112,114) nested relative to one another and overlapping in a thickness direction. The first and second portions (112,114) are adjacent to one another by sharing a common wall (110). The first portion (112) is provided on the suction side (88). The second portion (114)is provided on the pressure side (86).
摘要翻译： 用于燃气涡轮发动机的部件（64）包括提供压力和吸力侧（86,88）的外表面。 组件（64）中的冷却通道（90）包括蛇形通道（94），该蛇形通道具有分别被配置为提供彼此相反方向的流体流动的第一和第二道次（118,120）。 第一遍（118）包括相对于彼此嵌套并在厚度方向上重叠的第一和第二部分（112,114）。 第一和第二部分（112,114）通过共享公共壁（110）彼此相邻。 第一部分（112）设置在吸力侧（88）上。 第二部分（114）设置在压力侧（86）上。
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公开(公告)号：EP3196414B1
公开(公告)日：2018-11-28
申请号：EP17152351.7
申请日：2017-01-20
发明人： AUXIER, James Tilsley , ZELESKY, Mark F. , TELLER, Bret M. , JARIWALA, Parth , MONGILLO, Dominic J.
CPC分类号： F01D5/187 , F01D9/02 , F01D25/12 , F04D29/324 , F04D29/542 , F04D29/582 , F05D2220/32 , F05D2250/185 , F05D2260/202 , F05D2260/22141 , F05D2260/606 , Y02T50/676
摘要： An airfoil (201; 301; 401; 501) of a gas turbine engine is provided including a leading edge (212; 312) extending in a radial direction, a tip (232; 332; 432; 532) extending in an axial direction from the leading edge, a first rib (232; 332; 432; 532) extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity (220; 320; 420; 520) within the airfoil, a second rib (234; 334), the second rib and the first rib defining a serpentine cavity (228; 328; 428; 528) therein, a third rib (236; 336; 436; 536) extending axially within the tip, a flag tip cavity (226; 326; 426; 526) defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture (240; 340; 440; 540) formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.
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公开(公告)号：EP2956627B1
公开(公告)日：2018-07-25
申请号：EP14797049.5
申请日：2014-02-07
发明人： LEWIS, Scott D. , RAPP, Brandon M. , BEATTIE, Jeffrey S. , HOUGH, Matthew Andrew , TELLER, Bret M. , JACQUES, Jeffrey Michael , ASTERLIN, Max
CPC分类号： F01D5/186 , F01D5/18 , F01D9/023 , F01D9/065 , F01D11/08 , F01D25/12 , F05D2240/11 , F05D2240/35 , F05D2240/81 , F05D2260/202 , F23R3/002 , F23R3/02 , F23R3/04 , F23R3/06 , F23R2900/03041 , F23R2900/03042 , Y02T50/675 , Y02T50/676
摘要： A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.
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7.发明公开
公开(公告)号：EP2956627A2
公开(公告)日：2015-12-23
申请号：EP14797049.5
申请日：2014-02-07
发明人： LEWIS, Scott D. , RAPP, Brandon M. , BEATTIE, Jeffrey S. , HOUGH, Matthew Andrew , TELLER, Bret M. , JACQUES, Jeffrey Michael , ASTERLIN, Max
CPC分类号： F01D5/186 , F01D5/18 , F01D9/023 , F01D9/065 , F01D11/08 , F01D25/12 , F05D2240/11 , F05D2240/35 , F05D2240/81 , F05D2260/202 , F23R3/002 , F23R3/02 , F23R3/04 , F23R3/06 , F23R2900/03041 , F23R2900/03042 , Y02T50/675 , Y02T50/676
摘要： A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.
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公开(公告)号：EP3693546A1
公开(公告)日：2020-08-12
申请号：EP20156368.1
申请日：2020-02-10
发明人： NIEZELSKI, David A. , JENNINGS, Timothy J. , TELLER, Bret M. , MCBRIEN, John , SLAVENS, Thomas N.
摘要： Airfoils for gas turbine engines are described. The airfoils (320) include an airfoil body extending between a platform (348) and a tip (322), the airfoil body having a leading edge (322), a trailing edge (324), a pressure side (326), and a suction side (328), a serpentine cavity formed within the airfoil body and having an up-pass serpentine cavity (340), a down-pass serpentine cavity (342), and a trailing edge cavity (344), and a dead-end tip flag cavity (338) extending in a direction between the leading edge and the trailing edge, the dead-end tip flag cavity arranged between the serpentine cavity and the tip, wherein the dead-end tip flag cavity ends at a dead-end wall (366) located at a position between the leading edge and the trailing edge of the airfoil body. A corresponding core assembly for the formation of the airfoil is also provided.
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公开(公告)号：EP3241992B1
公开(公告)日：2019-01-30
申请号：EP17166119.2
申请日：2017-04-11
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公开(公告)号：EP3241992A1
公开(公告)日：2017-11-08
申请号：EP17166119.2
申请日：2017-04-11
CPC分类号： F01D5/186 , F01D5/18 , F01D9/02 , F01D11/08 , F01D25/12 , F05D2220/32 , F05D2230/21 , F05D2230/211 , F05D2240/122 , F05D2240/127 , F05D2240/304 , F05D2240/35 , F05D2260/202 , F05D2260/2212 , F23R3/002 , Y02T50/676
摘要： A casting core (130; 200) may comprise a tip flag cavity (134) having a forward pedestal (138, 139) and a first spear pedestal (136) disposed aft of the forward pedestal (138, 139). A trailing edge discharge cavity (132) may be separated from the tip flag cavity (134) and include a first row of pedestals (140). The first row of pedestals (140) may comprise a first racetrack pedestal (160). A second row of pedestals (142) may be disposed aft of the first row of pedestals (140) and include a second racetrack pedestal (166). A third row of pedestals (144) may be disposed aft of the second row of pedestals (142) and include a circular pedestal (168). A fourth row of pedestals (146) may be disposed aft of the third row of pedestals (144) and include a second spear pedestal (170).
摘要翻译： 铸造型芯（130; 200）可以包括具有前部基座（138,139）和设置在前部基座（138,139）后面的第一矛基座（136）的尖端旗腔（134）。 后缘排出腔（132）可以与末端旗腔（134）分开并且包括第一排基座（140）。 第一排基座（140）可以包括第一跑道基座（160）。 第二排基座（142）可以设置在第一排基座（140）的后面并且包括第二跑道基座（166）。 第三排基座（144）可以设置在第二排基座（142）的后面并且包括圆形基座（168）。 第四排基座（146）可以设置在第三排基座（144）的后面并且包括第二矛基座（170）。
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