公开(公告)号：US20150118034A1

公开(公告)日：2015-04-30

申请号：US14068070

申请日：2013-10-31

Applicant: Ching-Pang Lee ,  Benjamin E. Heneveld ,  Glenn E. Brown ,  Jill Klinger

Inventor： Ching-Pang Lee ,  Benjamin E. Heneveld ,  Glenn E. Brown ,  Jill Klinger

IPC: F01D5/18 ,  F01D9/02

CPC classification number: F01D5/187 ,  F01D9/02 ,  F05D2240/122 ,  F05D2240/304 ,  F05D2260/201 ,  F05D2260/2212

Abstract: A gas turbine engine component, including: a pressure side (12) having an interior surface (34); a suction side (14) having an interior surface (36); a trailing edge portion (30); and a plurality of suction side and pressure side impingement orifices (24) disposed in the trailing edge portion (30). Each suction side impingement orifice is configured to direct an impingement jet (48) at an acute angle (52) onto a target area (60) that encompasses a tip (140) of a chevron (122) within a chevron arrangement (120) formed in the suction side interior surface. Each pressure side impingement orifice is configured to direct an impingement jet at an acute angle onto an elongated target area that encompasses a tip of a chevron within a chevron arrangement formed in the pressure side interior surface.

Abstract translation: 一种燃气涡轮发动机部件，包括：具有内表面（34）的压力侧（12）; 吸入侧（14），其具有内表面（36）; 后缘部分（30）; 以及设置在后缘部分（30）中的多个吸力侧和压力侧冲击孔（24）。 每个吸入侧冲击孔构造成将锐角（52）的冲击射流（48）引导到目标区域（60）上，所述目标区域包围形成的人字形布置（120）内的人字形（122）的尖端（140） 在吸力侧内表面。 每个压力侧冲击孔构造成将冲击射流以锐角引导到细长的目标区域上，所述细长目标区域包含形成在压力侧内表面中的人字形布置中的人字形尖端。

公开(公告)号：US20130142666A1

公开(公告)日：2013-06-06

申请号：US13311630

申请日：2011-12-06

Applicant: Ching-Pang Lee ,  Glenn E. Brown ,  Benjamin E. Heneveld

Inventor： Ching-Pang Lee ,  Glenn E. Brown ,  Benjamin E. Heneveld

IPC: F01D5/18

CPC classification number: F01D5/187 ,  F01D5/186 ,  F05D2240/122 ,  F05D2240/304

Abstract: A turbine blade (10) including an airfoil (12) having multiple interior wall portions (70) each separating at least one chamber from another one of multiple chambers (46, 48, 50, 58, 60). In one embodiment a first wall portion (70-2) between first and second chambers (60, 52) includes first and second pluralities of flow paths (86P, 86S) extending through the first wall portion. The first wall portion includes a first region R1 having a first thickness, t, measurable as a distance between the chambers. One of the paths extends a first path distance, d, as measured from an associated path opening (78) in the first chamber (60), through the first region and to an exit opening (82) in the second chamber (52) which path distance is greater than the first thickness.

Abstract translation: 一种涡轮机叶片（10），包括具有多个内壁部分（70）的翼型件（12），每个内壁部分（70）将至少一个腔室与多个腔室（46,48,50,58,60）中的另一个分隔开。 在一个实施例中，第一和第二腔室（60,52）之间的第一壁部分（70-2）包括延伸穿过第一壁部分的第一和第二多个流动路径（86P，86S）。 第一壁部分包括具有第一厚度t的第一区域R1，其可测量为腔室之间的距离。 路径中的一个沿着第一腔室（60）中的相关联的路径开口（78）测量的第一路径距离（d）延伸穿过第一区域并延伸到第二腔室（52）中的出口开口（82） 路径距离大于第一厚度。

公开(公告)号：US09004866B2

公开(公告)日：2015-04-14

申请号：US13311630

申请日：2011-12-06

Applicant: Ching-Pang Lee ,  Glenn E. Brown ,  Benjamin E. Heneveld

Inventor： Ching-Pang Lee ,  Glenn E. Brown ,  Benjamin E. Heneveld

IPC: F01D5/18

CPC classification number: F01D5/187 ,  F01D5/186 ,  F05D2240/122 ,  F05D2240/304

Abstract: A turbine blade (10) including an airfoil (12) having multiple interior wall portions (70) each separating at least one chamber from another one of multiple chambers (46, 48, 50, 58, 60). In one embodiment a first wall portion (70-2) between first and second chambers (60, 52) includes first and second pluralities of flow paths (86P, 86S) extending through the first wall portion. The first wall portion includes a first region R1 having a first thickness, t, measurable as a distance between the chambers. One of the paths extends a first path distance, d, as measured from an associated path opening (78) in the first chamber (60), through the first region and to an exit opening (82) in the second chamber (52) which path distance is greater than the first thickness.

Abstract translation: 一种涡轮机叶片（10），包括具有多个内壁部分（70）的翼型件（12），每个内壁部分（70）将至少一个腔室与多个腔室（46,48,50,58,60）中的另一个分隔开。 在一个实施例中，第一和第二腔室（60,52）之间的第一壁部分（70-2）包括延伸穿过第一壁部分的第一和第二多个流动路径（86P，86S）。 第一壁部分包括具有第一厚度t的第一区域R1，其可测量为腔室之间的距离。 路径中的一个沿着第一腔室（60）中的相关联的路径开口（78）测量的第一路径距离（d）延伸穿过第一区域并延伸到第二腔室（52）中的出口开口（82） 路径距离大于第一厚度。

公开(公告)号：US09039371B2

公开(公告)日：2015-05-26

申请号：US14068070

申请日：2013-10-31

Applicant: Ching-Pang Lee ,  Benjamin E. Heneveld ,  Glenn E. Brown ,  Jill Klinger

Inventor： Ching-Pang Lee ,  Benjamin E. Heneveld ,  Glenn E. Brown ,  Jill Klinger

IPC: F01D5/18 ,  F01D9/02

CPC classification number: F01D5/187 ,  F01D9/02 ,  F05D2240/122 ,  F05D2240/304 ,  F05D2260/201 ,  F05D2260/2212

Abstract: A gas turbine engine component, including: a pressure side (12) having an interior surface (34); a suction side (14) having an interior surface (36); a trailing edge portion (30); and a plurality of suction side and pressure side impingement orifices (24) disposed in the trailing edge portion (30). Each suction side impingement orifice is configured to direct an impingement jet (48) at an acute angle (52) onto a target area (60) that encompasses a tip (140) of a chevron (122) within a chevron arrangement (120) formed in the suction side interior surface. Each pressure side impingement orifice is configured to direct an impingement jet at an acute angle onto an elongated target area that encompasses a tip of a chevron within a chevron arrangement formed in the pressure side interior surface.

Abstract translation: 一种燃气涡轮发动机部件，包括：具有内表面（34）的压力侧（12）; 吸入侧（14），其具有内表面（36）; 后缘部分（30）; 以及设置在后缘部分（30）中的多个吸力侧和压力侧冲击孔（24）。 每个吸入侧冲击孔构造成将锐角（52）的冲击射流（48）引导到目标区域（60）上，所述目标区域包围形成的人字形布置（120）内的人字形（122）的尖端（140） 在吸力侧内表面。 每个压力侧冲击孔构造成将冲击射流以锐角引导到细长的目标区域上，所述细长目标区域包含形成在压力侧内表面中的人字形布置中的人字形尖端。

公开(公告)号：US08951004B2

公开(公告)日：2015-02-10

申请号：US13657923

申请日：2012-10-23

Applicant: Ching-Pang Lee ,  Benjamin E. Heneveld

Inventor： Ching-Pang Lee ,  Benjamin E. Heneveld

IPC: F01D5/18

CPC classification number: F01D5/187 ,  F05D2250/185 ,  F05D2260/22141

Abstract: A cooling arrangement (82) for a gas turbine engine component, the cooling arrangement (82) having a plurality of rows (92, 94, 96) of airfoils (98), wherein adjacent airfoils (98) within a row (92, 94, 96) define segments (110, 130, 140) of cooling channels (90), and wherein outlets (114, 134) of the segments (110, 130) in one row (92, 94) align aerodynamically with inlets (132, 142) of segments (130, 140) in an adjacent row (94, 96) to define continuous cooling channels (90) with non continuous walls (116, 120), each cooling channel (90) comprising a serpentine shape.

Abstract translation: 一种用于燃气涡轮发动机部件的冷却装置（82），所述冷却装置（82）具有多排（92,94,96）的翼型件（98），其中在一排（92,94）内相邻的翼型件 ，96）限定冷却通道（90）的段（110,130,140），并且其中一排（92,94）中的段（110,130）的出口（114,134）在空气动力学上与入口（132， 142）在相邻排（94,96）中的段（130,140），以限定具有非连续壁（116,120）的连续冷却通道（90），每个冷却通道（90）包括蛇形形状。

公开(公告)号：US08764394B2

公开(公告)日：2014-07-01

申请号：US12985553

申请日：2011-01-06

Applicant: Ching-Pang Lee ,  John J. Marra ,  Gary B. Merrill ,  Benjamin E. Heneveld ,  Jill Klinger

Inventor： Ching-Pang Lee ,  John J. Marra ,  Gary B. Merrill ,  Benjamin E. Heneveld ,  Jill Klinger

IPC: F01D5/18 ,  F28F3/04 ,  F28F7/02

CPC classification number: F01D5/18 ,  F01D5/187 ,  F05D2240/304 ,  F05D2250/13 ,  F05D2260/2214 ,  F28F3/048 ,  F28F7/02

Abstract: A cooling channel (36, 36B) cools an exterior surface (40 or 42) or two opposed exterior surfaces (40 and 42). The channel has a near-wall inner surface (48, 50) with a width (W1). Interior side surfaces (52, 54) may converge to a reduced channel width (W2). The near-wall inner surface (48, 50) may have fins (44) aligned with a coolant flow (22). The fins may highest at mid-width of the near-wall inner surface. A two-sided cooling channel (36) may have two near-wall inner surfaces (48, 50) parallel to two respective exterior surfaces (40, 42), and may have an hourglass shaped transverse sectional profile. The tapered channel width (W1, W2) and the fin height profile (56A, 56B) increases cooling flow (22) into the corners (C) of the channel for more uniform and efficient cooling.

Abstract translation: 冷却通道（36,36B）冷却外表面（40或42）或两个相对的外表面（40和42）。 通道具有宽度（W1）的近壁内表面（48,50）。 内侧表面（52,54）可以会聚到减小的通道宽度（W2）。 近壁内表面（48,50）可以具有与冷却剂流（22）对准的翅片（44）。 翅片在近壁内表面的中间宽度处可能最高。 双面冷却通道（36）可以具有两个平行于两个相应的外表面（40,42）的近壁内表面（48,50），并且可以具有沙漏形横截面轮廓。 锥形通道宽度（W1，W2）和翅片高度轮廓（56A，56B）将冷却流（22）增加到通道的拐角（C），以实现更均匀和有效的冷却。

公开(公告)号：US08936067B2

公开(公告)日：2015-01-20

申请号：US13658045

申请日：2012-10-23

Applicant: Ching-Pang Lee ,  Benjamin E. Heneveld

Inventor： Ching-Pang Lee ,  Benjamin E. Heneveld

IPC: B22C9/10

CPC classification number: B22C9/10 ,  F01D5/186 ,  F01D5/187 ,  F05D2230/21 ,  F05D2230/211

Abstract: A ceramic casting core, including: a plurality of rows (162, 166, 168) of gaps (164), each gap (164) defining an airfoil shape; interstitial core material (172) that defines and separates adjacent gaps (164) in each row (162, 166, 168); and connecting core material (178) that connects adjacent rows (170, 174, 176) of interstitial core material (172). Ends of interstitial core material (172) in one row (170, 174, 176) align with ends of interstitial core material (172) in an adjacent row (170, 174, 176) to form a plurality of continuous and serpentine shaped structures each including interstitial core material (172) from at least two adjacent rows (170, 174, 176) and connecting core material (178).

Abstract translation: 一种陶瓷铸造芯，包括：多个间隔（164）的排（162,166,168），每个间隙（164）限定翼型形状; 间隙芯材料（172），其限定和分隔每排（162,166,168）中的相邻间隙（164）; 以及连接芯材料（178），其连接间隙芯材料（172）的相邻行（170,174,176）。 一排（170,174,176）中的间隙芯材（172）的端部与相邻排（170,174,176）中的间隙芯材（172）的端部对准，以形成多个连续和蛇形形状的结构 包括来自至少两个相邻行（170,174,176）和连接芯材（178）的间隙芯材料（172）。

公开(公告)号：US20120282108A1

公开(公告)日：2012-11-08

申请号：US13099521

申请日：2011-05-03

Applicant: Ching-Pang Lee ,  Shantanu P. Mhetras ,  Glenn E. Brown

Inventor： Ching-Pang Lee ,  Shantanu P. Mhetras ,  Glenn E. Brown

IPC: F01D5/18

CPC classification number: F01D5/20

Abstract: A squealer tip formed from a pressure side rib and a suction side rib extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side ribs may be positioned along the pressure side and the suction side of the turbine blade, respectively. The pressure and suction side ribs may include chamfered leading edges with film cooling holes having exhaust outlets positioned therein. The film cooling holes may be configured to be diffusion cooling holes with one or more tapered sections for reducing the velocity of cooling fluids and increasing the size of the convective surfaces.

Abstract translation: 公开了一种从压力侧肋和从涡轮叶片的尖端径向向外延伸的吸力侧肋形成的尖叫尖端。 压力和吸力侧肋可以分别沿涡轮叶片的压力侧和吸力侧定位。 压力和吸力侧肋可以包括倒角的前缘，其中膜冷却孔具有位于其中的排气口。 膜冷却孔可以被配置为具有一个或多个锥形部分的扩散冷却孔，用于降低冷却流体的速度并增加对流表面的尺寸。

公开(公告)号：US08684691B2

公开(公告)日：2014-04-01

申请号：US13099521

申请日：2011-05-03

Applicant: Ching-Pang Lee ,  Shantanu P. Mhetras ,  Glenn E. Brown

Inventor： Ching-Pang Lee ,  Shantanu P. Mhetras ,  Glenn E. Brown

IPC: F01D5/20

CPC classification number: F01D5/20

Abstract: A squealer tip formed from a pressure side rib and a suction side rib extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side ribs may be positioned along the pressure side and the suction side of the turbine blade, respectively. The pressure and suction side ribs may include chamfered leading edges with film cooling holes having exhaust outlets positioned therein. The film cooling holes may be configured to be diffusion cooling holes with one or more tapered sections for reducing the velocity of cooling fluids and increasing the size of the convective surfaces.

Abstract translation: 公开了一种从压力侧肋和从涡轮叶片的尖端径向向外延伸的吸力侧肋形成的尖叫尖端。 压力和吸力侧肋可以分别沿涡轮叶片的压力侧和吸力侧定位。 压力和吸力侧肋可以包括倒角的前缘，其中膜冷却孔具有位于其中的排气口。 膜冷却孔可以被配置为具有一个或多个锥形部分的扩散冷却孔，用于降低冷却流体的速度并增加对流表面的尺寸。

公开(公告)号：US09630277B2

公开(公告)日：2017-04-25

申请号：US12723880

申请日：2010-03-15

Applicant: Ching-Pang Lee ,  Mrinal Munshi ,  Gm S. Azad ,  Jae Y. Um

Inventor： Ching-Pang Lee ,  Mrinal Munshi ,  Gm S. Azad ,  Jae Y. Um

IPC: F02C7/12 ,  B23P6/00 ,  F01D5/00 ,  F01D5/14 ,  F01D5/18

CPC classification number: B23P6/007 ,  F01D5/005 ,  F01D5/145 ,  F01D5/186 ,  F05D2230/30 ,  F05D2230/80 ,  F05D2240/81 ,  Y10T29/49318

Abstract: A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform.