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公开(公告)号：US10301966B2
公开(公告)日：2019-05-28
申请号：US14957731
申请日：2015-12-03
Applicant: United Technologies Corporation
Inventor： Mohammed Ennacer , Russell J. Bergman , Jason B. Moran
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. At least one of the inner and outer platforms includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1 for the inner platform or Table 2 for the outer platform. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).
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公开(公告)号：US10280780B2
公开(公告)日：2019-05-07
申请号：US14928147
申请日：2015-10-30
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman
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 state. 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.
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公开(公告)号：US09670790B2
公开(公告)日：2017-06-06
申请号：US13659969
申请日：2012-10-25
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman , Leonard A. Bach
CPC classification number: F01D9/042 , F05D2230/60 , F05D2230/64 , F05D2240/12 , Y10T29/49323
Abstract: A turbine vane for a gas turbine engine has an inner platform, an outer platform, at least one airfoil extending between the inner and outer platforms, and a tab radially extending inward from a front side of the inner platform. The tab contains a mounting aperture and an identification aperture that identifies an engine in which the turbine vane may be installed.
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公开(公告)号：US20160208629A1
公开(公告)日：2016-07-21
申请号：US14965501
申请日：2015-12-10
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman , Thurman Carlo Dabbs
CPC classification number: F01D9/041 , F01D11/08 , F01D21/003 , F01D25/246 , F05D2220/32 , F05D2240/12 , F05D2240/80 , F05D2260/30
Abstract: A vane for a gas turbine engine includes a radially inner platform and a radially outer platform. At least two airfoils extend between the radially inner platform and the radially outer platform. At least two anti-rotation tabs extend axially from the radially outer platform.
Abstract translation: 用于燃气涡轮发动机的叶片包括径向内部平台和径向外部平台。 至少两个翼型件在径向内部平台和径向外部平台之间延伸。 至少两个防旋转翼片从径向外部平台轴向延伸。
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公开(公告)号：US20160160654A1
公开(公告)日：2016-06-09
申请号：US14953745
申请日：2015-11-30
Applicant: United Technologies Corporation
Inventor： Mohammed Ennacer , Russell J. Bergman , Jason B. Moran
CPC classification number: F01D25/12 , F01D9/041 , F05D2240/12 , F05D2240/81 , 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. At least one of the inner and outer platforms includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1 for the inner platform or Table 2 for the outer platform. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. 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。 内平台和外平台中的至少一个包括膜冷却孔，其具有基本符合表1中针对内平台或外平台的表2所示的笛卡尔坐标的外部突出点。 相对于零坐标，笛卡尔坐标由轴向坐标，圆周坐标和径向坐标提供。 薄膜冷却孔相对于0.20英寸（5.0mm）的指定坐标具有径向表面公差。
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Patent Agency Ranking
公开(公告)号：US20160153304A1
公开(公告)日：2016-06-02
申请号：US14942484
申请日：2015-11-16
Applicant: United Technologies Corporation
Inventor： Andrew S. Aggarwala , Russell J. Bergman , Mark A. Boeke , Kyle C. Lana , Daniel A. Snyder
CPC classification number: F01D11/006 , F01D5/22 , F01D11/001 , F01D11/02
Abstract: An airfoil stage of a turbine engine includes an upstream airfoil assembly, a downstream airfoil assembly in rotational relationship to the upstream airfoil assembly and a rim seal assembly integrated therebetween. The rim seal assembly may include a sloped downstream portion of a platform of the upstream airfoil assembly, an upstream segment of a platform of the downstream airfoil assembly and a nub that projects radially outward from the upstream segment. The downstream portion and the upstream segment are spaced from one-another defining a cooling cavity therebetween for the flow of cooling air. The portion and segment overlap axially such that the nub is axially aligned to the downstream portion for improved cooling effectiveness and a reduction of core airflow into the cooling cavity.
Abstract translation: 涡轮发动机的翼型段包括上游翼型组件，与上游翼型组件成旋转关系的下游翼型组件和集成在其间的边缘密封组件。 轮辋密封组件可以包括上游翼型组件的平台的倾斜下游部分，下游翼型组件的平台的上游部分和从上游部分径向向外突出的凸块。 下游部分和上游段彼此间隔开，在其间限定冷却空气的流动。 该部分和部分轴向重叠，使得该凸块与下游部分轴向对齐，以改善冷却效果并减少进入冷却腔的芯气流。
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公开(公告)号：US20140147263A1
公开(公告)日：2014-05-29
申请号：US13659969
申请日：2012-10-25
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman , Leonard A. Bach
IPC: F01D9/04
CPC classification number: F01D9/042 , F05D2230/60 , F05D2230/64 , F05D2240/12 , Y10T29/49323
Abstract: A turbine vane for a gas turbine engine has an inner platform, an outer platform, at least one airfoil extending between the inner and outer platforms, and a tab radially extending inward from a front side of the inner platform. The tab contains a mounting aperture and an identification aperture that identifies an engine in which the turbine vane may be installed.
Abstract translation: 用于燃气涡轮发动机的涡轮叶片具有内平台，外平台，在内平台和外平台之间延伸的至少一个翼型，以及从内平台的前侧向内径向延伸的突片。 标签包含安装孔和识别孔，其识别可安装涡轮叶片的发动机。
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公开(公告)号：US11377965B2
公开(公告)日：2022-07-05
申请号：US15450087
申请日：2017-03-06
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Shawn J. Gregg , Dominic J. Mongillo , Michael Leslie Clyde Papple , Russell J. Bergman , Mohammed Ennacer
Abstract: A component for a gas turbine engine includes, among other things, an airfoil that extends between a leading edge and a trailing edge and a cooling circuit disposed inside of the airfoil. The cooling circuit includes at least one core cavity that extends inside of the airfoil, a baffle received within the at least one core cavity, a plurality of pedestals positioned adjacent to the at least one core cavity and a first plurality of axial ribs positioned between the plurality of pedestals and the trailing edge of the airfoil.
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39.发明申请公开(公告)号：US20190309636A1
公开(公告)日：2019-10-10
申请号：US15947254
申请日：2018-04-06
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman , Alan W. Stoner
IPC: F01D5/22
Abstract: A blade for a gas turbine engine includes an airfoil that includes an internal passage. A shroud is arranged at an end of the airfoil and has a shroud perimeter. Axially spaced knife edges extend radially from the shroud. An area is provided between the knife edges. A pocket is recessed into the area and is circumscribed by a perimeter edge that is arranged interiorly of the shroud perimeter. An outlet fluidly connects the internal passage to the pocket.
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公开(公告)号：US20170335700A1
公开(公告)日：2017-11-23
申请号：US15159935
申请日：2016-05-20
Applicant: United Technologies Corporation
Inventor： Russell J. Bergman , Brett Alan Bartling
CPC classification number: F01D9/041 , F01D5/18 , F01D9/065 , F05D2220/32 , F05D2230/10 , F05D2230/232 , F05D2230/237 , F05D2240/10 , F05D2240/121 , F05D2240/81 , F05D2250/324 , F05D2260/202
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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