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公开(公告)号：US20180216472A1
公开(公告)日：2018-08-02
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：US20170145827A1
公开(公告)日：2017-05-25
申请号：US15018025
申请日：2016-02-08
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis , Kyle C. Lana , Jason David Liles
CPC classification number: F01D5/145 , F01D5/141 , F01D5/20 , F01D5/22 , F02C3/04 , F05D2220/32 , F05D2240/303 , F05D2240/304 , F05D2240/305 , F05D2240/306 , F05D2240/307 , F05D2250/711 , Y02T50/671 , Y02T50/673
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 squealer pocket located in a surface of the tip, wherein the squealer pocket is at least partially surrounded by a first surface of a wall located between the squealer pocket and a pressure side of the airfoil, wherein the first surface of the wall has a convex configuration 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：US20160230564A1
公开(公告)日：2016-08-11
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.
Abstract translation: 根据本公开的示例的涡轮叶片包括平台，翼型末端以及平台和翼型顶端之间的翼型部分。 机翼部分具有与翼型件末端径向间隔开的空腔和在空腔和翼型端之间径向放置的多个冷却通道。 多个冷却通道中的每一个限定了与翼型件末端相邻的出口。 多个冷却通道内的每个内部特征被配置成计量到出口的流量。
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公开(公告)号：US10753210B2
公开(公告)日：2020-08-25
申请号：US15968865
申请日：2018-05-02
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 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 shielding sidewall cavity located adjacent one of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, a second shielding sidewall cavity located adjacent the other of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, and a shielded sidewall cavity located between the first shielding sidewall cavity and the second shielding sidewall cavity, wherein the shielded sidewall cavity is not adjacent either of the pressure or suction sides proximate the root and transitions to be proximate at least one of the pressure and suction sides proximate the tip.
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公开(公告)号：US10513931B2
公开(公告)日：2019-12-24
申请号：US14974418
申请日：2015-12-18
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Scott D. Lewis , Kyle C. Lana
Abstract: An airfoil is provided. The airfoil comprises a crossover and an impingement cavity in fluid communication with the crossover and having an internal surface. At least a portion of the internal surface comprises an undulating internal surface. A plurality of trip strips may be disposed on the at least a portion of the internal surface to define the undulating internal surface. A gas turbine engine and an internally-cooled engine part are also provided.
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Patent Agency Ranking
公开(公告)号：US20190338652A1
公开(公告)日：2019-11-07
申请号：US15968865
申请日：2018-05-02
Applicant: United Technologies Corporation
Inventor： Timothy J. Jennings , Tracy A. Propheter-Hinckley , Kyle C. Lana
IPC: F01D5/18
Abstract: Airfoils for gas turbine engines are provided. The airfoils include an airfoil 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 shielding sidewall cavity located adjacent one of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, a second shielding sidewall cavity located adjacent the other of the pressure and suction sides proximate the root of the airfoil body and extending radially toward the tip, and a shielded sidewall cavity located between the first shielding sidewall cavity and the second shielding sidewall cavity, wherein the shielded sidewall cavity is not adjacent either of the pressure or suction sides proximate the root and transitions to be proximate at least one of the pressure and suction sides proximate the tip.
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公开(公告)号：US20190338648A1
公开(公告)日：2019-11-07
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：US10253635B2
公开(公告)日：2019-04-09
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
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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公开(公告)号：US20180306058A1
公开(公告)日：2018-10-25
申请号：US15496977
申请日：2017-04-25
Applicant: United Technologies Corporation
Inventor： Scott D Lewis , Kyle C. Lana
CPC classification number: F01D25/12 , F01D5/082 , F01D5/3007 , F01D11/006 , F01D11/04 , F05D2220/32 , F05D2240/81 , F05D2260/202
Abstract: An airfoil may include an airfoil body, a root and a platform disposed between the airfoil body and the root. The platform may have a first mating surface and a second mating surface. The platform may include a pocket defined by an inner diameter surface of the platform proximate the first mating surface. A channel may be defined in the platform with an outlet of the channel defined in the second mating surface.
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公开(公告)号：US20170350255A1
公开(公告)日：2017-12-07
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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