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公开(公告)号：US10907482B2
公开(公告)日：2021-02-02
申请号：US16126482
申请日：2018-09-10
Applicant: United Technologies Corporation
Inventor： Daniel A. Snyder , Jaimie Schweitzer , Jonathan Perry Sandoval , Edwin Otero , Kyle C. Lana , Scott D. Virkler
Abstract: An example turbine includes a turbine disk and a turbine blade. The turbine disk includes a plurality of lugs and a plurality of slots. Each of the plurality of lugs is located between two of the plurality of slots. Each of the plurality of lugs includes a tab that extends radially outwardly from an end of the lug. The tab has a first section having an upper surface and a second section having an upper surface. The upper surface of the first section is inclined greater than the upper surface of the second section, and the upper surface of the first section of the tab defines a contour. The turbine blade includes a root received in one of the plurality of slots and a platform, and the platform has a lower surface defining a contour. When the turbine blade is received in one of the plurality of slots, the lower surface of the platform is located above a portion of the upper surface of the first section of the tab, and the contour of the upper surface of the first section of the tab matches the contour of the lower surface of the platform.
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2.发明申请公开(公告)号：US20160177755A1
公开(公告)日：2016-06-23
申请号：US14862351
申请日：2015-09-23
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Steven D. Porter , John T. Ols , Kyle C. Lana
CPC classification number: F01D9/02 , F01D1/04 , F01D5/081 , F01D11/001 , F05D2240/80 , Y02T50/676
Abstract: A gas-turbine engine is provided. The gas-turbine engine comprises a high pressure turbine with an aft blade platform. A static structure may be disposed aft of the high pressure turbine and proximate a cavity defined by the aft blade platform. A vane of the static structure may have a vane platform with a shaped tip extending into the cavity.
Abstract translation: 提供燃气涡轮发动机。 燃气涡轮发动机包括具有后叶片平台的高压涡轮机。 静态结构可以设置在高压涡轮机的后方并且靠近由后叶片平台限定的空腔。 静态结构的叶片可以具有叶片平台，其具有延伸到空腔中的成形尖端。
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公开(公告)号：US10815788B2
公开(公告)日：2020-10-27
申请号：US15419612
申请日：2017-01-30
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis , Kyle C. Lana
Abstract: A component for a gas turbine engine, the component having: a cooling slot located on a surface of the component, the cooling slot being defined by a plurality of diffuser portions each extending from a respective one of a plurality of cooling openings providing cooling fluid to the cooling slot.
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公开(公告)号：US10557354B2
公开(公告)日：2020-02-11
申请号：US14914084
申请日：2014-08-13
Applicant: United Technologies Corporation
Inventor： Edwin Otero , Kyle C. Lana
Abstract: A gas turbine engine component includes spaced apart walls that provide a cooling passage that extends in a first direction. A cross-over rib joins the walls and extends along the first direction. The cross-over rib has holes that extend in a second direction transverse to the first direction. A row of at least one pedestal joins the walls and extends along the first direction. The row and the cross-over rib overlap one another in the second direction.
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公开(公告)号：US20160153865A1
公开(公告)日：2016-06-02
申请号：US14812723
申请日：2015-07-29
Applicant: United Technologies Corporation
Inventor： Kyle C. Lana , Stephen D. Haddock
CPC classification number: G01M15/14 , F01D5/186 , F01D21/003 , F05D2260/80 , G01B11/00 , Y02T50/676
Abstract: A method of measuring untwist in an gas turbine engine airfoil, the method includes measuring first and second cooling holes on an exterior airfoil surface of an airfoil, determining untwist in a chord-wise direction and/or creep in a radial direction based upon the measurement, evaluating the determination relative to reference information for the airfoil, and outputting a part status based upon the evaluation.
Abstract translation: 一种在燃气涡轮发动机翼型中测量解卷积的方法，该方法包括测量翼型件的外部翼型表面上的第一和第二冷却孔，基于测量确定沿着方向取向和/或径向蠕动 评估相对于翼型件的参考信息的确定，并且基于评估输出部件状态。
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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：US10941663B2
公开(公告)日：2021-03-09
申请号：US15972637
申请日：2018-05-07
Applicant: United Technologies Corporation
Inventor： Timothy J. Jennings , Tracy A. Propheter-Hinckley , Kyle C. Lana
Abstract: Airfoils for gas turbine engines are provided. The airfoils include a body extending between leading and trailing edges in an axial direction, between pressure and suction sides in a circumferential direction, and between a root and tip in a radial direction. A first transitioning leading edge cavity is located adjacent one of the sides proximate the root of the body and transitions axially toward the leading edge as the first transitioning leading edge cavity extends radially toward the tip. A second transitioning leading edge cavity is adjacent the other side and adjacent the leading edge proximate the root of the body and transitions axially toward the trailing edge as the second transitioning leading edge cavity extends radially toward the tip. A portion of the second transitioning leading edge cavity shields a portion of the first transitioning leading edge cavity proximate the root of the body.
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公开(公告)号：US10801331B2
公开(公告)日：2020-10-13
申请号：US15175108
申请日：2016-06-07
Applicant: United Technologies Corporation
Inventor： Patrick D. Couture , Scott D. Lewis , Kyle C. Lana
Abstract: A gas turbine engine blade includes a blade portion having a leading edge and a trailing edge. A first surface connects the leading edge to the trailing edge and a second surface connects the leading edge to the trailing edge. A tip section is located at a first end of the blade portion and includes a pocket protruding into the tip section from an outermost end of the tip section. The pocket has a first side wall adjacent the first surface and a second side wall adjacent the second surface. At least one of the first side wall and the second side wall have a curve distinct from a curve of the corresponding adjacent surface.
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公开(公告)号：US20190010810A1
公开(公告)日：2019-01-10
申请号：US16126482
申请日：2018-09-10
Applicant: United Technologies Corporation
Inventor： Daniel A. Snyder , Jaimie Schweitzer , Jonathan Perry Sandoval , Edwin Otero , Kyle C. Lana , Scott D. Virkler
Abstract: An example turbine includes a turbine disk and a turbine blade. The turbine disk includes a plurality of lugs and a plurality of slots. Each of the plurality of lugs is located between two of the plurality of slots. Each of the plurality of lugs includes a tab that extends radially outwardly from an end of the lug. The tab has a first section having an upper surface and a second section having an upper surface. The upper surface of the first section is inclined greater than the upper surface of the second section, and the upper surface of the first section of the tab defines a contour. The turbine blade includes a root received in one of the plurality of slots and a platform, and the platform has a lower surface defining a contour. When the turbine blade is received in one of the plurality of slots, the lower surface of the platform is located above a portion of the upper surface of the first section of the tab, and the contour of the upper surface of the first section of the tab matches the contour of the lower surface of the platform.
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公开(公告)号：US20180245470A1
公开(公告)日：2018-08-30
申请号：US15966049
申请日：2018-04-30
Applicant: United Technologies Corporation
Inventor： Dominic J. Mongillo, JR. , Jeffrey R. Levine , Kyle C. Lana , Sasha M. Moore
CPC classification number: F01D5/186 , F01D5/187 , F01D5/20 , F05D2240/307 , F05D2260/2214 , F05D2260/22141 , Y02T50/673 , Y02T50/676
Abstract: A turbine blade according to an example of the present disclosure includes, among other things, a platform extending from a root section, an airfoil section extending radially from the platform to an airfoil tip, a plurality of cooling passages defined in an external wall of the airfoil tip, the plurality of cooling passages extending radially between the airfoil tip and a cavity in the airfoil section bounded by the external wall, and each of the plurality of cooling passages defining an inlet port along the cavity and an exit port adjacent the airfoil tip, and at least one internal feature within each of the plurality of cooling passages that meter flow to the respective exit port.
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