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    TAPERED THERMAL COATING FOR AIRFOIL
    1.
    发明申请
    TAPERED THERMAL COATING FOR AIRFOIL 审中-公开
    用于气流的热转印热涂层

    公开(公告)号：WO2013172903A2

    公开(公告)日：2013-11-21

    申请号：PCT/US2013/027232

    申请日：2013-02-22

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： LEVINE, Jeffrey R. ,  GREGG, Shawn J.

    IPC: F01D5/28 ,  F01D5/14 ,  F02C7/00 ,  F02K3/00

    CPC classification number: F01D5/188 ,  Y02T50/676

    Abstract: An airfoil comprises pressure and suction surfaces extending axially from a leading edge to a trailing edge and radially from a root section to a tip section, the root section and the tip section defining a span therebetween. A thermal coating extends from the root section of the airfoil toward the tip section of the airfoil. A relative coating thickness of the thermal coating decreases by at least thirty percent at full span in the tip section, as compared to minimum span in the root section.

    Abstract translation: 翼型件包括从前缘到后缘轴向延伸的压力和抽吸表面，并且从根部到尖端部分径向地延伸，根部和末端部分限定了它们之间的跨度。 热涂层从翼型的根部朝向翼型的末端部分延伸。 与根部中的最小跨度相比，热涂层的相对涂层厚度在末端部分的全跨度处降低至少30％。

    IMPROVED COOLING FOR A TURBINE AIRFOIL TRAILING EDGE
    2.
    发明申请
    IMPROVED COOLING FOR A TURBINE AIRFOIL TRAILING EDGE 审中-公开

    公开(公告)号：WO2014007889A3

    公开(公告)日：2014-01-09

    申请号：PCT/US2013/034220

    申请日：2013-03-28

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： LEVINE, Jeffrey R. ,  MONGILLO, Dominic J. ,  DONNELL, Brandon S.

    IPC: F01D5/18 ,  F02C7/12

    Abstract: An assembly for a gas turbine engine includes a first platform and an airfoil extending from the first platform. The airfoil includes a first fillet, pressure side biased discharge openings, and a first center cooling discharge opening. A pressure side wall of the airfoil and the first platform form an acute angle at the trailing edge. The first fillet is formed around a perimeter of the airfoil where the airfoil extends from the first platform. The pressure side biased cooling discharge openings are along the trailing edge outside of the first fillet. Each pressure side biased cooling discharge opening extends from the trailing edge along the pressure side wall. The first center cooling discharge opening extends along the trailing edge into the first fillet and is centrally located between the pressure side wall and the suction side wall.

    GAS TURBINE ENGINE SERPENTINE COOLING PASSAGE
    5.
    发明申请
    GAS TURBINE ENGINE SERPENTINE COOLING PASSAGE 审中-公开
    燃气涡轮发动机SERPENTINE冷却通风

    公开(公告)号：WO2014042955A1

    公开(公告)日：2014-03-20

    申请号：PCT/US2013/058282

    申请日：2013-09-05

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： PEREZ, Rafel A. ,  PIETRASZKIEWICZ, Edward F. ,  LEVINE, Jeffrey R. ,  MONGILLO JR., Dominic J.

    IPC: F01D5/08 ,  F02C7/12 ,  F02K3/00

    CPC classification number: F01D5/188 ,  F01D5/187 ,  F05D2250/183 ,  F05D2250/185 ,  F05D2260/2212 ,  F05D2260/22141 ,  F05D2270/17 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A gas turbine engine component includes a structure having a cooling passage providing upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend. The downstream portion includes an outer wall opposite the inner wall to provide a downstream region extending between the inner and outer walls. A turbulence promoter extends from the outer wall adjacent to the bend in the downstream portion. The turbulence promoter is absent from a stagnation region adjoining the inner wall adjacent to the bend in the downstream portion

    Abstract translation: 燃气涡轮发动机部件包括具有冷却通道的结构，该冷却通道提供通过内壁彼此分离并通过弯曲部流体连接的上游和下游部分。 下游部分包括与内壁相对的外壁，以提供在内壁和外壁之间延伸的下游区域。 湍流促进剂从与下游部分中的弯曲部相邻的外壁延伸。 在与下游部分中的弯曲部相邻的内壁的停滞区域中不存在湍流促进剂

    BLADE TIP COOLING ARRANGEMENT
    8.
    发明公开
    BLADE TIP COOLING ARRANGEMENT 有权
    SCHAUFELSPITZENKÜHLUNGSANORDNUNG

    公开(公告)号：EP3081753A1

    公开(公告)日：2016-10-19

    申请号：EP16155312.8

    申请日：2016-02-11

    Applicant: United Technologies Corporation

    Inventor： MONGILLO, Dominic J., Jr. ,  LEVINE, Jeffrey R. ,  LANA, Kyle C. ,  MOORE, Sasha M.

    IPC: F01D5/18 ,  F01D9/02 ,  F01D11/08

    CPC classification number: F01D5/186 ,  F01D5/187 ,  F01D5/20 ,  F05D2240/307 ,  F05D2260/2214 ,  F05D2260/22141 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A turbine blade (161) includes a platform (162), an airfoil tip (164), and an airfoil section (165) between the platform (162) and the airfoil tip (164). The airfoil section (165) has a cavity (177) spaced radially from the airfoil tip (164) and a plurality of cooling passages (182) radially between the cavity (177) and the airfoil tip (164). Each of the plurality of cooling passages (182) defines an exit port (184) adjacent the airfoil tip (164). An internal feature (188) within each of the plurality of cooling passages (182) is configured to meter flow to the exit port (184).

    Abstract translation: 涡轮叶片（161）包括平台（162），翼型件尖端（164）和平台（162）与翼型件（164）之间的翼型部分（165）。 机翼部分（165）具有与翼型件（164）径向间隔开的空腔（177）和在空腔（177）和翼型件（164）之间径向的多个冷却通道（182）。 多个冷却通道（182）中的每一个限定与翼型件尖端（164）相邻的出口（184）。 多个冷却通道（182）内部的内部特征（188）被配置成流量到出口（184）。

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