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公开(公告)号：US20190101021A1
公开(公告)日：2019-04-04
申请号：US15723459
申请日：2017-10-03
Applicant: United Technologies Corporation
Inventor： Carey Clum , Dominic J. Mongillo, JR.
Abstract: A component for a gas turbine engine, includes an external surface bounding a hot gas path of the gas turbine engine, and a cooling passage configured to deliver a cooling airflow therethrough. The cooling passage includes an internal surface located opposite the external surface, together defining a component wall. A plurality of trip strip features are located along the internal surface having a trip strip height extending from the internal surface and a trip strip width extending along the internal surface in a flow direction of the cooling airflow through the cooling passage. A ratio of a trip strip pitch between adjacent trip strip features in a width direction and the trip strip height is less than 5. One or more cooling film bleed holes extend from the internal surface to the external surface and are located between adjacent trip strip features of the plurality of trip strip features.
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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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公开(公告)号：US20180223676A1
公开(公告)日：2018-08-09
申请号：US15426082
申请日：2017-02-07
Applicant: United Technologies Corporation
Inventor： Brandon W. Spangler , Dominic J. Mongillo, JR.
IPC: F01D5/18
CPC classification number: F01D5/187 , F01D5/189 , F01D9/02 , F01D9/065 , F01D25/12 , F05D2220/32 , F05D2230/21 , F05D2240/125 , F05D2240/126 , F05D2240/81 , F05D2250/185 , F05D2260/221
Abstract: Turn caps for airfoils of gas turbine engines including cavity sidewalls, a first turn cap divider extending between the cavity sidewalls and defining a turning cavity between the first turn cap divider and the cavity sidewalls, and a second turn cap divider disposed radially inward within the turning cavity. A first turning path is defined between the first turn cap divider and the second turn cap divider and a second turning path is defined radially inward of the second turn cap divider and a merging chamber is formed in the turn cap wherein fluid flows through the first turning path and the second turning path are merged, the merging chamber, the first turning path, and the second turning path forming the turning cavity.
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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：US20160032730A1
公开(公告)日：2016-02-04
申请号：US14774157
申请日：2014-03-13
Applicant: UNITED TECHNOLOGIES CORPORATION
Inventor： Seth J. Thomen , Edward F. Pietraszkiewicz , Dominic J. Mongillo, JR. , Wieslaw A. Chlus , Erik R. Granstrand
Abstract: An airfoil includes a cooling air passage for receiving a cooling air flow. A chevron including a first rib and a second rib extends from a common tip is disposed within the cooling passage for generating a turbulent flow to improve heat transfer. The chevron includes an angle between the first rib and the second rib that is greater than 90 degrees.
Abstract translation: 翼型件包括用于接收冷却空气流的冷却空气通道。 包括第一肋和第二肋的人字形从共同的尖端延伸设置在冷却通道内，用于产生湍流以改善热传递。 人字形包括大于90度的第一肋和第二肋之间的角度。
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Patent Agency Ranking
公开(公告)号：US20190226345A1
公开(公告)日：2019-07-25
申请号：US16371239
申请日：2019-04-01
Applicant: United Technologies Corporation
Inventor： Brandon W. Spangler , Dominic J. Mongillo, JR.
Abstract: A gas turbine engine component with a core includes a first core portion configured to provide a first cooling passage. A second core portion is spaced from the first core portion and configured to provide a second cooling passage. The second core portion includes a longitudinal leg and a resupply leg transverse to and intersecting the longitudinal leg. The resupply leg has a terminal end and is configured to provide a resupply channel A connector interconnects the terminal end to the first core portion. The connector is configured to provide a resupply hole.
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公开(公告)号：US20190226343A1
公开(公告)日：2019-07-25
申请号：US15876579
申请日：2018-01-22
Applicant: United Technologies Corporation
Inventor： Nicholas M. LoRicco , Ricardo Trindade , Dominic J. Mongillo, JR.
Abstract: Components for gas turbine engines are provided. The components include a hot external wall that is exposed to hot gaspath air when installed within a gas turbine engine, and an interior impingement wall, wherein the interior impingement wall defines a feed cavity and at least one impingement cavity is defined between the impingement wall and the external wall. The impingement wall includes a plurality of impingement holes that fluidly connect the feed cavity to the at least one impingement cavity, the external wall includes a plurality of film holes that fluidly connect the at least one impingement cavity to an exterior surface of the external wall, and wherein the only source of cooling air within the at least one impingement cavity is the feed cavity.
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18.发明申请公开(公告)号：US20190211690A1
公开(公告)日：2019-07-11
申请号：US15886054
申请日：2018-02-01
Applicant: United Technologies Corporation
Inventor： Nicholas M. LoRicco , Dominic J. Mongillo, JR.
CPC classification number: F01D9/065 , F05D2240/81 , F05D2250/75 , F05D2260/202 , F05D2260/204 , F05D2260/22141
Abstract: An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform. An airfoil skin cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The airfoil skin cooling passage extends to a platform skin cooling passage arranged in the platform. A resupply hole fluidly connects the main-body core cooling passage and at least one of the airfoil skin cooling passage and the platform skin cooling passage.
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公开(公告)号：US20190195074A1
公开(公告)日：2019-06-27
申请号：US15852400
申请日：2017-12-22
Applicant: United Technologies Corporation
Inventor： Dominic J. Mongillo, JR. , Allan N. Arisi
Abstract: Components for gas turbine engines are provided. The components include a hybrid skin core cooling cavity defined by a cold wall and a hot wall, wherein the hot wall is exposed to an exterior environment of the component, a hybrid resupply hole formed in the cold wall and fluidly connecting a cold cavity and the hybrid skin core cooling cavity, and a resupply cover located on the cold wall and within the hybrid skin core cooling cavity and positioned relative to the hybrid resupply hole to shield resupply air injected from the cold cavity into the hybrid skin core cooling cavity to minimize losses as the resupply air mixes with air flowing within the hybrid skin core cooling cavity.
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公开(公告)号：US20190169993A1
公开(公告)日：2019-06-06
申请号：US15832407
申请日：2017-12-05
Applicant: United Technologies Corporation
Inventor： Christopher Whitfield , Carey Clum , Dominic J. Mongillo, JR.
CPC classification number: F01D5/187 , F02C3/04 , F02C9/18 , F05D2220/32 , F05D2240/35 , F05D2260/201 , F05D2260/202 , F05D2260/204 , F05D2260/221
Abstract: An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole cluster includes first and second resupply holes that each fluidly interconnect the core and skin passages. The first and second resupply holes intersect one another and terminate in an exit at the skin passage.
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