-
21.发明申请公开(公告)号：US20170002663A1
公开(公告)日：2017-01-05
申请号：US14790231
申请日：2015-07-02
Applicant: United Technologies Corporation
Inventor： San Quach , Scott D. Lewis , Steven Bruce Gautschi , Christopher King , Christopher Corcoran , David Donald Chapdelaine
IPC: F01D5/18
CPC classification number: F01D5/20 , F05D2260/202 , Y02T50/673 , Y02T50/676
Abstract: A gas turbine engine airfoil includes a body that provides an exterior airfoil surface that extends in a radial direction to a tip. The exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction. The tip includes a squealer pocket that has a recess surface. A cooling passage is arranged in the body. Each of the cooling holes extends from an inlet at the cooling passage to an outlet at the recessed surface. The inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface. The angular direction is toward at least one of the forward and aft directions.
Abstract translation: 燃气涡轮发动机翼型件包括提供沿径向方向延伸到尖端的外部翼面表面的主体。 外表面具有向前方向的前缘和向后方的后缘。 尖端包括具有凹陷表面的尖叫口袋。 冷却通道设置在体内。 每个冷却孔从冷却通道的入口延伸到凹陷表面处的出口。 入口和出口相对于凹入表面以角度方向成一定角度布置。 角度方向朝着前后方向中的至少一个方向。
-
公开(公告)号：US11306919B2
公开(公告)日：2022-04-19
申请号：US16224012
申请日：2018-12-18
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis
Abstract: A cooling hole arrangement for a combustor panel including a first group of cooling holes, each of the first cooling holes extending from a first inlet to a first outlet. Also included is a second group of cooling holes, each of the second cooling holes extending from a second inlet to a second outlet, each of the first cooling holes and the second cooling holes spaced from an edge of the combustor panel in an axial direction to be arranged substantially parallel to the edge of the combustor panel. At least a portion of the first inlet of each first cooling holes is axially overlapped with a portion of the second inlet of the second cooling holes. The first outlet of each of the first cooling holes is closer to the edge of the combustor panel than the second outlet of the second cooling holes is to the edge.
-
公开(公告)号：US11156117B2
公开(公告)日：2021-10-26
申请号：US15137044
申请日：2016-04-25
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis
Abstract: A seal for a gas turbine engine includes a plurality of seal arc segments. Each of the seal arc segments includes radially inner and outer sides and sloped first and second circumferential sides. The seal arc segments are circumferentially arranged about an axis such that the sloped first and second circumferential sides define gaps circumferentially between adjacent ones of the seal arc segments. Each of the gaps extends from the radially inner sides along a respective central gap axis that slopes with respect to a radial direction from the axis.
-
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.
-
公开(公告)号：US10273811B2
公开(公告)日：2019-04-30
申请号：US14707092
申请日：2015-05-08
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis
Abstract: A component for use in a turbine engine includes a fore edge connected to an aft edge via a first surface and a second surface. Multiple cooling passages are defined within the turbine engine component. A first skin core passage is defined immediately adjacent one of the first surface and the second surface. At least 80% of coolant entering the first skin core passage is expelled from the turbine engine component at the aft edge.
-
Patent Agency Ranking
公开(公告)号：US20180230811A1
公开(公告)日：2018-08-16
申请号：US15355173
申请日：2016-11-18
Inventor： Scott D. Lewis , Stephen Lynch , Shane Haydt
IPC: F01D5/18
CPC classification number: F01D5/186 , F05D2230/12 , F05D2250/19 , F05D2250/312 , F05D2250/314 , F05D2260/202 , Y02T50/676
Abstract: A component for a gas turbine engine including a body having at least one internal cooling cavity and a plurality of film cooling holes disposed along a first edge of the body. At least one of the film cooling holes includes a metering section defining an axis, and a diffuser section having a centerline. The centerline of the diffuser section is offset from the axis of the metering section.
-
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.
-
公开(公告)号：US09982542B2
公开(公告)日：2018-05-29
申请号：US14802459
申请日：2015-07-17
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis , Christopher Corcoran
IPC: F01D5/18
CPC classification number: F01D5/187 , F01D5/186 , F05D2240/81 , F05D2260/201 , Y02T50/676
Abstract: An airfoil structure for a gas turbine engine includes an airfoil which includes a leading edge and a trailing edge. A platform is located adjacent a first end of the airfoil and includes a core passage that extends through the platform, a mate-face for engaging an adjacent airfoil structure and a set of impingement cooling holes in communication with the core passage that extend through the mate-face adjacent the trialing edge of the airfoil.
-
公开(公告)号：US20170306781A1
公开(公告)日：2017-10-26
申请号：US15137044
申请日：2016-04-25
Applicant: United Technologies Corporation
Inventor： Scott D. Lewis
CPC classification number: F01D11/08 , F01D5/02 , F01D5/12 , F01D11/005 , F01D11/04 , F01D11/24 , F01D25/12 , F05D2220/32 , F05D2240/11
Abstract: A seal for a gas turbine engine includes a plurality of seal arc segments. Each of the seal arc segments includes radially inner and outer sides and sloped first and second circumferential sides. The seal arc segments are circumferentially arranged about an axis such that the sloped first and second circumferential sides define gaps circumferentially between adjacent ones of the seal arc segments. Each of the gaps extends from the radially inner sides along a respective central gap axis that slopes with respect to a radial direction from the axis.
-
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.
-
-
-
-
-
-
-
-
-
Search Results
63
records
(took 0.021 seconds)
