公开(公告)号：US20200080424A1

公开(公告)日：2020-03-12

申请号：US16581895

申请日：2019-09-25

Applicant: United Technologies Corporation

Inventor： Matthew A. Devore ,  Matthew S. Gleiner

IPC: F01D5/18 ,  F01D9/02 ,  F01D25/12

Abstract: A gas turbine engine component includes an airfoil that has a leading edge and extends from a first side of a platform that has a leading edge overhang. A root portion extends from a second side of the platform opposite the first side. A cooling passage extends through the platform and beneath a trailing edge of the airfoil. The cooling passage includes an inlet located on a second opposite side of the platform. The inlet is located axially upstream of the leading edge of the airfoil and the inlet is located axially upstream of the root portion in the leading edge overhang.

公开(公告)号：US10253636B2

公开(公告)日：2019-04-09

申请号：US14997992

申请日：2016-01-18

Applicant: United Technologies Corporation

Inventor： Matthew A. Devore ,  Eleanor D. Kaufman

IPC: F01D5/18 ,  F01D25/12 ,  F01D9/02 ,  F01D9/06

Abstract: A baffle insert for a component of a gas turbine engine is provided. The baffle insert having: a first fluid conduit having a first interior cavity extending therethrough; a second fluid conduit having a second interior cavity extending therethrough; and a member located between the first fluid conduit and the second fluid conduit, wherein the member fluidly couples the first interior cavity to an exterior of the second fluid conduit, and wherein the member fluidly couples the second interior cavity to an exterior of the first fluid conduit and wherein the first interior cavity is isolated from the second interior cavity.

公开(公告)号：US20180298770A1

公开(公告)日：2018-10-18

申请号：US15490299

申请日：2017-04-18

Applicant: United Technologies Corporation

Inventor： Matthew A. Devore ,  Jonathan Ortiz

IPC: F01D9/06 ,  F01D9/04 ,  F02C7/18

Abstract: Turbines comprising a first stator section having a plurality of first vanes, a first rotating section having a plurality of first blades, a second stator section having a plurality of second vanes, a “primary TOBI assembly” having an “aft-facing, forward-positioned TOBI” configured to direct an airflow from the first stator section in an aftward direction toward the first rotating section, the primary TOBI assembly supplying high pressure cooling flow to leading edges of the first blades of the first rotating section, and a “secondary TOBI assembly” having a “forward-facing, aft-positioned TOBI” configured to direct an airflow from the second stator section in a forward direction toward the first rotating section, the secondary TOBI assembly supplying low pressure cooling flow to non-leading edge portions of the first blades of the first rotating section.

公开(公告)号：US20180291760A1

公开(公告)日：2018-10-11

申请号：US15484166

申请日：2017-04-11

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Jonathan Ortiz ,  Matthew A. Devore ,  Lane Mikal Thornton ,  James D. Hill

IPC: F01D11/20 ,  F02C3/04 ,  F02C9/18 ,  F01D5/02

Abstract: A gas turbine engine comprises a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface defining a diffuser chamber radially outwardly of the combustor, and a cooling air chamber wall positioned outwardly of the diffuser chamber and the combustor, and radially inwardly of a second wall to define a cooling air chamber. The turbine section includes a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap taps air having been compressed by the compressor and is passed through a heat exchanger. Air downstream of the heat exchanger passes into the cooling chamber, and then to the blade outer air seal.

公开(公告)号：US20180209286A1

公开(公告)日：2018-07-26

申请号：US15416394

申请日：2017-01-26

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Brandon W. Spangler ,  Carey Clum ,  Matthew A. Devore

IPC: F01D9/06 ,  F02C3/04 ,  F02C7/12

CPC classification number: F01D9/065 ,  F01D5/187 ,  F02C7/18 ,  F05D2220/32 ,  F05D2240/122 ,  F05D2260/20 ,  F05D2260/202 ,  F05D2260/2212 ,  Y02T50/676

Abstract: A gas turbine engine component comprises a body having a leading edge, a trailing edge, and a radial span. One internal channel in the body provides an upstream supply pressure. Another internal channel in body receives the upstream supply pressure and provides a downstream supply pressure. At least one axial rib separates an internal area adjacent to the trailing edge into a plurality of individual cavities. At least one pressure regulating feature is located at an entrance to at least one individual cavity entrance to control downstream supply pressure to the trailing edge. Exits formed in the trailing edge communicate with an exit pressure. The rib and pressure regulating features cooperate such that the downstream supply pressure mimics the exit pressure along the radial span. A method of manufacturing a gas turbine engine component and a method of controlling flow in a gas turbine engine component are also disclosed.

公开(公告)号：US20180135437A1

公开(公告)日：2018-05-17

申请号：US15354055

申请日：2016-11-17

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Tracy A. Propheter-Hinckley ,  Michael G. McCaffrey ,  Raymond Surace ,  Matthew A. Devore

IPC: F01D5/28 ,  F01D5/18 ,  F01D9/04 ,  F01D25/12 ,  F04D29/38 ,  F04D29/54 ,  B23P15/04

CPC classification number: F01D5/284 ,  B23P15/04 ,  B32B18/00 ,  C04B2235/6028 ,  C04B2237/38 ,  C04B2237/62 ,  C04B2237/84 ,  F01D5/147 ,  F01D5/187 ,  F01D5/282 ,  F01D9/041 ,  F01D9/042 ,  F01D9/065 ,  F01D25/12 ,  F04D29/388 ,  F04D29/542 ,  F05D2220/32 ,  F05D2230/60 ,  F05D2240/121 ,  F05D2240/303 ,  F05D2260/201 ,  F05D2260/202 ,  F05D2260/204 ,  F05D2300/6033 ,  Y02T50/672 ,  Y02T50/673 ,  Y02T50/676

Abstract: A method of fabricating a ceramic turbine engine article includes building a wall of the article from preceramic layers, wherein the building includes arranging the preceramic layers around one or more sacrificial core elements, converting the preceramic layers to ceramic, and removing the one or more sacrificial core elements to leave one or more cavities in the wall.

公开(公告)号：US20170175539A1

公开(公告)日：2017-06-22

申请号：US14976041

申请日：2015-12-21

Applicant: United Technologies Corporation

Inventor： Benjamin F. Hagan ,  Matthew A. Devore ,  Dominic J. Mongillo ,  Ryan Alan Waite

IPC: F01D5/18 ,  F02C3/04 ,  F01D25/24 ,  F01D9/04 ,  F01D25/12

CPC classification number: F01D5/187 ,  F01D9/041 ,  F01D9/065 ,  F01D25/12 ,  F01D25/24 ,  F02C3/04 ,  F05D2220/32 ,  F05D2240/121 ,  F05D2240/303 ,  F05D2240/35 ,  F05D2250/232 ,  F05D2250/323 ,  F05D2260/201 ,  F05D2260/202 ,  F05D2260/205 ,  Y02T50/676

Abstract: A flowpath component for a gas turbine engine includes a leading edge, a trailing edge connected to the leading edge via a first surface and a second surface, an impingement cavity internal to the flowpath component, the impingement cavity being aligned with one of the leading edge and the trailing edge, a cooling passage extending at least partially through the flowpath component, and a plurality of crossover holes connecting the cooling passage to the impingement cavity. At least one of the crossover holes is aligned normal to an expected direction of fluid flow through the cooling passage and is unaligned with an axial line drawn perpendicular to a stacking line of the flowpath component.

公开(公告)号：US20160326883A1

公开(公告)日：2016-11-10

申请号：US15107493

申请日：2015-01-02

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： Lane Thornton ,  Matthew A. Devore ,  Dominic J. Mongillo ,  Steven Bruce Gautschi

IPC: F01D5/18

Abstract: A gas turbine engine component comprises an airfoil with a suction side and pressure side extending from a leading edge to a trailing edge. There are a plurality of cooling holes adjacent the leading edge, with the cooling holes having a non-circular shape, with a longer dimension and a smaller dimension. The airfoil defines a radial direction from a radially outer end to a radially inner end, and radially outer of the cooling holes spaced toward the radially outer end, which have the longer dimension extending closer to parallel to the radial direction. Radially inner cooling holes closer to the radially inner end having the longer dimension extend to be closer to perpendicular relative to the radial direction compared to the radially outer cooling holes.

Abstract translation: 燃气涡轮发动机部件包括具有从前缘延伸到后缘的吸力侧和压力侧的翼型件。 在前缘附近有多个冷却孔，冷却孔具有非圆形形状，具有较长的尺寸和较小的尺寸。 翼型件限定从径向外端到径向内端的径向方向，并且冷却孔的径向外部与径向外端间隔开，其具有更靠近平行于径向的较长尺寸。 与径向外侧冷却孔相比，更靠近具有较长尺寸的径向内端的径向内部冷却孔相对于径向更接近垂直。

公开(公告)号：US20160076382A1

公开(公告)日：2016-03-17

申请号：US14821831

申请日：2015-08-10

Applicant: United Technologies Corporation

Inventor： Matthew A. Devore ,  Matthew S. Gleiner ,  Douglas C. Jenne

IPC: F01D5/18 ,  B22C9/10 ,  F01D25/12 ,  F01D9/04 ,  F01D11/08

CPC classification number: F01D5/186 ,  B22C9/10 ,  B22C9/106 ,  F01D5/187 ,  F01D9/041 ,  F01D11/08 ,  F01D25/12 ,  F05D2220/32 ,  F05D2230/211 ,  F05D2240/12 ,  F05D2240/126 ,  F05D2240/30 ,  F05D2240/55 ,  F05D2240/81 ,  F05D2260/202 ,  Y02T50/676

Abstract: A gas turbine engine component according to an exemplary aspect of the present disclosure includes, among other things, at least one cooling passage. The at least one cooling passage includes a first wall and a second wall bounding the at least one cooling passage, the first wall having a plurality of first surface features and the second wall having a plurality of second surface features. The plurality of first surface features and the plurality of second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the plurality of first surface features and the plurality of second surface features. The plurality of first surface features has a first profile, and the plurality of second surface features has a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed.

Abstract translation: 根据本公开的示例性方面的燃气涡轮发动机部件尤其包括至少一个冷却通道。 所述至少一个冷却通道包括限定所述至少一个冷却通道的第一壁和第二壁，所述第一壁具有多个第一表面特征，所述第二壁具有多个第二表面特征。 多个第一表面特征和多个第二表面特征被布置成使得冷却通道的宽度沿着由多个第一表面特征和多个第二表面特征限定的冷却通道的长度而变化。 多个第一表面特征具有第一轮廓，并且多个第二表面特征具有第二不同轮廓。 还公开了一种用于在飞行器部件中形成冷却通道的铸芯。

公开(公告)号：US20140090384A1

公开(公告)日：2014-04-03

申请号：US13629705

申请日：2012-09-28

Applicant: UNITED TECHNOLOGIES CORPORATION

Inventor： John McBrien ,  Brandon W. Spangler ,  Matthew A. Devore

IPC: F01D5/18 ,  F23R3/02 ,  F01D9/02

CPC classification number: F01D5/186 ,  F05D2250/14 ,  F05D2250/141 ,  F05D2250/314 ,  F23R3/002 ,  F23R3/06 ,  F23R2900/00018 ,  F23R2900/03042 ,  Y02T50/676

Abstract: A gas turbine engine component includes a structure having an exterior surface. A cooling hole extends from a cooling passage to the exterior surface to provide an exit area on the exterior surface that is substantially circular in shape. A gas turbine engine includes a compressor section and a turbine section. A combustor is provided between the compressor and turbine sections. A component in at least one of the compressor and turbine sections has an exterior surface. A film cooling hole extends from a cooling passage to the exterior surface to provide an exit area that is substantially circular in shape. A method of machining a film cooling hole includes providing a component having an internal cooling passage and an exterior surface, machining a film cooling hole from the exterior surface to the internal cooling passage to provide a substantially circular exit area on the exterior surface.

Abstract translation: 燃气涡轮发动机部件包括具有外表面的结构。 冷却孔从冷却通道延伸到外表面，以在外表面上形成大致圆形形状的出口区域。 燃气涡轮发动机包括压缩机部和涡轮部。 在压缩机和涡轮机部分之间提供燃烧器。 压缩机和涡轮机部分中的至少一个中的部件具有外表面。 薄膜冷却孔从冷却通道延伸到外表面以提供基本上圆形的出口区域。 一种加工薄膜冷却孔的方法包括提供具有内部冷却通道和外部表面的部件，从外表面到内部冷却通道加工薄膜冷却孔，以在外部表面上提供基本上圆形的出口区域。