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公开(公告)号：US11021976B2
公开(公告)日：2021-06-01
申请号：US14862351
申请日：2015-09-23
Applicant: United Technologies Corporation
Inventor： Steven D. Porter , John T. Ols , Kyle C. Lana
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.
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公开(公告)号：US10907479B2
公开(公告)日：2021-02-02
申请号：US15972429
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 an airfoil body extending between a leading edge and a trailing edge in an axial direction, between a pressure side and a suction side in a circumferential direction, and between a root and a tip in a radial direction, a first transitioning leading edge cavity located proximate the leading edge proximate the root of the airfoil body and transitioning axially toward the trailing edge as the first transitioning leading edge cavity extends radially toward the tip, and a second transitioning leading edge cavity located aft of the first transitioning leading edge cavity proximate the root of the airfoil body and transitioning axially toward the leading edge as the second transitioning leading edge cavity extends radially toward the tip. The second transitioning leading edge cavity includes an impingement sub-cavity and a film sub-cavity along the leading edge and proximate the tip.
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公开(公告)号：US10677066B2
公开(公告)日：2020-06-09
申请号：US15018072
申请日：2016-02-08
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis , Kyle C. Lana
Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a tip shelf located in a surface of the tip proximate to a pressure side of the airfoil, wherein the tip shelf has a ledge portion extending from the pressure side to a wall portion extending upwardly from the ledge portion to the tip and wherein the wall portion is configured to have a convex portion with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.
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公开(公告)号：US10648353B2
公开(公告)日：2020-05-12
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
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.
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公开(公告)号：US20190338649A1
公开(公告)日：2019-11-07
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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Patent Agency Ranking
公开(公告)号：US09995147B2
公开(公告)日：2018-06-12
申请号：US14619343
申请日：2015-02-11
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, an airfoil tip, and an airfoil section between the platform and the airfoil tip. The airfoil section has a cavity spaced radially from the airfoil tip and a plurality of cooling passages radially between the cavity and the airfoil tip. Each of the plurality of cooling passages defines an exit port adjacent the airfoil tip. An internal feature within each of the plurality of cooling passages is configured to meter flow to the exit port.
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公开(公告)号：US09810075B2
公开(公告)日：2017-11-07
申请号：US14663752
申请日：2015-03-20
Applicant: United Technologies Corporation
Inventor： Kyle C. Lana , John E. Paul
CPC classification number: F01D5/26 , F01D5/10 , F01D5/147 , F01D5/3007 , F01D11/006 , F01D11/008 , F01D25/06 , F05D2220/32 , F05D2240/24 , F05D2240/55 , F05D2260/96 , F05D2300/10 , Y02T50/673
Abstract: A damper seal received in a cavity of a turbine blade includes a central body with two ends and a two sides. A portion extends from each end, one side has a faceted edge, and a projection extends from the other side to receive a lug of the turbine blade to align the damper seal relative to the turbine blade.
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公开(公告)号：US20170145828A1
公开(公告)日：2017-05-25
申请号：US15018072
申请日：2016-02-08
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis , Kyle C. Lana
Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a tip shelf located in a surface of the tip proximate to a pressure side of the airfoil, wherein the tip shelf has a ledge portion extending from the pressure side to a wall portion extending upwardly from the ledge portion to the tip and wherein the wall portion is configured to have a convex portion with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.
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公开(公告)号：US20160273367A1
公开(公告)日：2016-09-22
申请号：US14663752
申请日：2015-03-20
Applicant: United Technologies Corporation
Inventor： Kyle C. Lana , John E. Paul
CPC classification number: F01D5/26 , F01D5/10 , F01D5/147 , F01D5/3007 , F01D11/006 , F01D11/008 , F01D25/06 , F05D2220/32 , F05D2240/24 , F05D2240/55 , F05D2260/96 , F05D2300/10 , Y02T50/673
Abstract: A damper seal received in a cavity of a turbine blade includes a central body with two ends and a two sides. A portion extends from each end, one side has a faceted edge, and a projection extends from the other side to receive a lug of the turbine blade to align the damper seal relative to the turbine blade.
Abstract translation: 容纳在涡轮叶片的空腔中的阻尼器密封件包括具有两端和两侧的中心体。 一部分从每个端部延伸，一侧具有小平面边缘，并且突出部从另一侧延伸以接收涡轮叶片的凸耳，以相对于涡轮叶片对准阻尼器密封件。
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30.发明申请GAS TURBINE ENGINE AIRFOIL CROSSOVER AND PEDESTAL RIB COOLING ARRANGEMENT 审中-公开气体涡轮发动机空气涡轮机和PEDESTAL RIB冷却装置公开(公告)号：US20160201477A1
公开(公告)日：2016-07-14
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.
Abstract translation: 燃气涡轮发动机部件包括间隔开的壁，其提供沿第一方向延伸的冷却通道。 交叉肋连接壁并沿第一方向延伸。 交叉肋具有沿与第一方向横切的第二方向延伸的孔。 至少一个基座的一排连接壁并沿第一方向延伸。 行和交叉肋在第二方向上彼此重叠。
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