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    VIBRATION DAMPERS FOR TURBINE BLADES
    2.
    发明申请
    VIBRATION DAMPERS FOR TURBINE BLADES 审中-公开
    涡轮叶片的振动阻尼器

    公开(公告)号：WO2015026416A2

    公开(公告)日：2015-02-26

    申请号：PCT/US2014/039710

    申请日：2014-05-28

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： KNOWLTON, Robert ,  ASTERLIN, Max ,  RAPP, Brandon, M. ,  JARIWALA, Parth

    IPC: F01D5/16

    CPC classification number: F01D5/10 ,  F01D5/081 ,  F01D5/18 ,  F01D5/26 ,  F01D11/006 ,  F05D2240/57 ,  F05D2260/96 ,  Y02T50/676

    Abstract: A turbine blade vibration damper is described. The damper has an axially-extending main body defining an inner surface and opposed damping surface. The main body also has an edge spanning from the inner surface to the damping surface and slotted apertures extending through the main body. The slotted apertures extend between the inner surface and damping surface. Adjacent pairs of slotted apertures have elongated shapes defining longitudinal axes that intersect at a point off the surface of the damper.

    Abstract translation: 描述了一种涡轮叶片减振器。 阻尼器具有限定内表面和相对的阻尼表面的轴向延伸主体。 主体还具有从内表面延伸到阻尼表面的边缘和延伸穿过主体的开槽孔。 开槽孔在内表面和阻尼表面之间延伸。 相邻的一对开槽孔具有细长的形状，从而形成纵向轴线，该纵向轴线在阻尼器表面的一点处相交。

    GAS TURBINE ENGINE AIRFOIL COOLING PASSAGE CONFIGURATION
    3.
    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

    Applicant: United Technologies Corporation

    Inventor： AUXIER, James T. ,  JARIWALA, Parth ,  ZELESKY, Mark F. ,  TELLER, Bret M.

    IPC: F01D5/18

    CPC classification number: F01D5/186 ,  F01D5/187 ,  F05D2250/185 ,  Y02T50/676

    Abstract: 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).

    Abstract translation: 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。

    INTERNALLY COOLED AIRFOIL
    8.
    发明公开
    INTERNALLY COOLED AIRFOIL 审中-公开
    内部冷却的机翼

    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

    Applicant: United Technologies Corporation

    Inventor： AUXIER, James T. ,  JARIWALA, Parth ,  TELLER, Bret M.

    IPC: F01D5/18 ,  F01D25/12

    CPC classification number: 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

    Abstract: 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).

    Abstract translation: 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）。

    DUAL-FED AIRFOIL TIP
    9.
    发明公开
    DUAL-FED AIRFOIL TIP 有权
    双翼飞机翼尖

    公开(公告)号：EP3196414A1

    公开(公告)日：2017-07-26

    申请号：EP17152351.7

    申请日：2017-01-20

    Applicant: United Technologies Corporation

    Inventor： AUXIER, James Tilsley ,  ZELESKY, Mark F. ,  TELLER, Bret M. ,  JARIWALA, Parth ,  MONGILLO, Dominic J.

    IPC: F01D5/18 ,  F01D5/30 ,  F02C3/16

    CPC classification number: F01D5/187 ,  F01D9/02 ,  F01D25/12 ,  F04D29/324 ,  F04D29/542 ,  F04D29/582 ,  F05D2220/32 ,  F05D2250/185 ,  F05D2260/202 ,  F05D2260/22141 ,  F05D2260/606 ,  Y02T50/676

    Abstract: 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.

    Abstract translation: 提供了一种燃气涡轮发动机的翼型件（201; 301; 401; 501），该翼型件包括在径向方向上延伸的前缘（212; 312），沿轴向方向延伸的尖端（232; 332; 432; 532） 所述前缘，在所述翼型中径向延伸的第一肋（232; 332; 432; 532），所述前缘和所述第一肋在所述翼型内限定出前缘腔（220; 320; 420; 520）;第二肋 （234; 334），所述第二肋和所述第一肋在其中限定了蛇形腔（228; 328; 428; 528），在所述尖端内轴向延伸的第三肋（236; 336; 436; 536） 由所述第三肋，所述前缘和所述尖端限定的所述前缘腔（226; 326; 426; 526），所述前缘腔流体地连接至所述旗尖端腔;以及旁通孔（240; 340; 440; 540） 所述第一肋和所述第三肋，所述旁路孔配置成将所述蛇形腔与所述旗尖端腔流体连接。

    VIBRATION DAMPERS FOR TURBINE BLADES
    10.
    发明公开
    VIBRATION DAMPERS FOR TURBINE BLADES 审中-公开
    涡轮叶片的振动阻尼器

    公开(公告)号：EP3004557A2

    公开(公告)日：2016-04-13

    申请号：EP14837891.2

    申请日：2014-05-28

    Applicant: United Technologies Corporation

    Inventor： KNOWLTON, Robert ,  ASTERLIN, Max ,  RAPP, Brandon, M. ,  JARIWALA, Parth

    IPC: F01D5/16 ,  F01D5/10

    CPC classification number: F01D5/10 ,  F01D5/081 ,  F01D5/18 ,  F01D5/26 ,  F01D11/006 ,  F05D2240/57 ,  F05D2260/96 ,  Y02T50/676

    Abstract: A turbine blade vibration damper is described. The damper has an axially-extending main body defining an inner surface and opposed damping surface. The main body also has an edge spanning from the inner surface to the damping surface and slotted apertures extending through the main body. The slotted apertures extend between the inner surface and damping surface. Adjacent pairs of slotted apertures have elongated shapes defining longitudinal axes that intersect at a point off the surface of the damper.

    Abstract translation: 描述了涡轮叶片减振器。 阻尼器具有限定内表面和相对的阻尼表面的轴向延伸主体。 主体还具有从内表面延伸到阻尼表面的边缘和延伸穿过主体的开槽孔。 开槽孔在内表面和阻尼表面之间延伸。 相邻的一对槽孔具有细长的形状，其限定纵向轴线，所述纵向轴线在阻尼器的表面处的点处相交。

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