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    TURBINE AIRFOIL SEGMENT HAVING FILM COOLING HOLE ARRANGEMENT
    25.
    发明申请
    TURBINE AIRFOIL SEGMENT HAVING FILM COOLING HOLE ARRANGEMENT 审中-公开
    涡轮机空气分段带有薄膜冷却孔安排

    公开(公告)号：US20160194966A1

    公开(公告)日：2016-07-07

    申请号：US14957686

    申请日：2015-12-03

    Applicant: United Technologies Corporation

    Inventor： Russell J. Bergman ,  Jason B. Moran ,  Mohammed Ennacer

    IPC: F01D5/18 ,  F02C7/18 ,  F02C3/04 ,  F01D9/04 ,  F01D25/12

    CPC classification number: F01D25/12 ,  F01D5/186 ,  F01D9/041 ,  F01D9/065 ,  F05D2250/74 ,  F05D2260/202 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676

    Abstract: A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. The airfoil includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in one of Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).

    Abstract translation: 涡轮机翼段包括通过至少一个翼型件连接的内平台和外平台。 翼型件包括通过间隔开的第一和第二侧连接以提供外部翼面表面的前缘和后缘。 翼型件包括具有大致符合表1和表2中所示的笛卡尔坐标的外部突出点的膜冷却孔。笛卡尔坐标由轴向坐标，圆周坐标和径向坐标提供， 相对于零坐标，并且膜冷却孔相对于0.20英寸（5.0mm）的指定坐标具有径向表面公差。

    INTERNALLY COOLED TURBINE PLATFORM
    26.
    发明申请
    INTERNALLY COOLED TURBINE PLATFORM 审中-公开
    内部冷却涡轮平台

    公开(公告)号：US20160146044A1

    公开(公告)日：2016-05-26

    申请号：US14945556

    申请日：2015-11-19

    Applicant: United Technologies Corporation

    Inventor： Russell J. Bergman

    IPC: F01D25/12 ,  F01D9/04 ,  F01D11/08 ,  F02C3/04

    CPC classification number: F01D25/12 ,  F01D9/04 ,  F01D9/041 ,  F01D11/08 ,  F01D25/246 ,  F02C3/04 ,  F05D2220/32 ,  F05D2240/12 ,  F05D2240/24 ,  F05D2240/35 ,  F05D2240/55 ,  F05D2260/201 ,  Y02T50/675

    Abstract: A turbine flowpath component for a gas turbine engine includes a platform defining at least one internal cooling cavity and a turbine flowpath component cover disposed radially outward of the platform. The turbine flowpath component cover defines a radially outward edge of the at least one internal cooling cavity. The turbine flowpath component cover further defines a transition region in which a radial depth of the at least one internal cooling cavity varies from a first depth to a second depth.

    Abstract translation: 用于燃气涡轮发动机的涡轮流路部件包括限定至少一个内部冷却空腔的平台和设置在该平台的径向外侧的涡轮流路部件盖。 涡轮流路部件盖限定至少一个内部冷却腔的径向向外的边缘。 涡轮流动路径部件盖进一步限定过渡区域，其中至少一个内部冷却腔的径向深度从第一深度变化到第二深度。

    SEALING SYSTEMS
    27.
    发明申请
    SEALING SYSTEMS 审中-公开
    密封系统

    公开(公告)号：US20160123170A1

    公开(公告)日：2016-05-05

    申请号：US14928147

    申请日：2015-10-30

    Applicant: United Technologies Corporation

    Inventor： Russell J. Bergman

    IPC: F01D11/00 ,  F01D9/04 ,  F16J15/02

    CPC classification number: F01D11/005 ,  F01D9/02 ,  F01D9/041 ,  F05D2220/32 ,  F05D2230/10 ,  F05D2230/60 ,  F05D2240/12 ,  F05D2240/55 ,  F05D2240/80 ,  F16J15/022

    Abstract: A sealing system for a turbomachine includes an annular seal defining a centerline axis and an inner diameter sealing surface. The sealing system also includes a turbomachine component radially inboard of the annular seal having a surface defined on an outer diameter surface. The sealing surface of the turbomachine component has a radius of curvature configured to be non-concentric with respect to the inner diameter surface of the annular seal in a cold sate. The radius of curvature of the surface is configured to be substantially concentric with respect to the inner diameter surface of the annular seal in a hot state.

    Abstract translation: 用于涡轮机的密封系统包括限定中心线轴线和内径密封表面的环形密封件。 密封系统还包括在环形密封件的径向内侧的涡轮机部件，其具有限定在外径表面上的表面。 涡轮机组件的密封表面具有相对于冷态中的环形密封件的内径表面非同心的曲率半径。 表面的曲率半径被构造成在热状态下相对于环形密封件的内径表面基本上是同心的。

    Multi-Airfoil Split and Rejoin Method
    28.
    发明申请
    Multi-Airfoil Split and Rejoin Method 审中-公开
    多翼分裂和重合法

    公开(公告)号：US20150369070A1

    公开(公告)日：2015-12-24

    申请号：US14762630

    申请日：2013-12-16

    Applicant: UNITED TECHNOLOGIES CORPORATION

    Inventor： Edward R. Bareiss ,  Russell J. Bergman ,  Bradley T. Duelm ,  Michael L. Miller ,  Paul M. Pellet

    IPC: F01D9/04 ,  F01D25/12

    Abstract: A method for working an airfoil cluster is disclosed. The method may include attaching a first datum to a first portion of the airfoil cluster, and joining a second portion of the airfoil cluster to the first portion, the second portion having a second datum substantially aligned with the first datum in a common plane spaced away from the first and second portions.

    Abstract translation: 公开了一种用于加工翼型组的方法。 该方法可以包括将第一基准附接到翼型群的第一部分，以及将翼型群的第二部分连接到第一部分，第二部分具有基本上与第一基准对准的公共平面间隔开的第二基准 从第一和第二部分。

    REPAIR OF CORE POSITIONING FEATURES IN CAST COMPONENTS
    29.
    发明申请
    REPAIR OF CORE POSITIONING FEATURES IN CAST COMPONENTS 有权

    公开(公告)号：US20210245237A1

    公开(公告)日：2021-08-12

    申请号：US16786291

    申请日：2020-02-10

    Applicant: United Technologies Corporation

    Inventor： Russell J. Bergman ,  Carl Busta ,  Lea Dynnette Castle

    IPC: B22C9/06 ,  B22C21/14 ,  B22C9/10

    Abstract: An investment casting process for manufacturing a cast component is provided. The investment casting process includes forming a core, casting the cast component about the core such that a core positioning feature provides a location of an anticipated pilot hole in the cast component, removing the core from the cast component once the casting is completed, locating, forming and sizing a pilot hole to form a resized pilot hole that can receive a sealing plug and installing the sealing plug into the resized pilot hole.

    Internal cooling of stator vanes
    30.
    发明授权
    Internal cooling of stator vanes 有权

    公开(公告)号：US10352182B2

    公开(公告)日：2019-07-16

    申请号：US15159935

    申请日：2016-05-20

    Applicant: United Technologies Corporation

    Inventor： Russell J. Bergman ,  Brett Alan Bartling

    IPC: F01D5/18 ,  F01D9/04 ,  F01D9/06

    Abstract: A stator for a gas turbine engine includes a stator vane, a first cooling passage located at the stator to provide a cooling fluid flow to a first portion of the stator, and a second cooling passage located at the stator to provide a cooling fluid flow to a second portion of the stator. A connection passage extends at least partially through the stator to connect a first cooling passage inlet of the first cooling passage to a second cooling passage inlet of the second cooling passage. The cooling fluid flow is directed from a common cooling flow source into the first cooling passage and the second cooling passage via the first cooling passage inlet.

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