公开(公告)号：US5498133A

公开(公告)日：1996-03-12

申请号：US471291

申请日：1995-06-06

Applicant: Ching-Pang Lee

Inventor： Ching-Pang Lee

IPC: F01D5/18

CPC classification number: F01D5/186 ,  F05D2240/305 ,  F05D2240/306 ,  F05D2250/52 ,  F05D2260/202

Abstract: A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges and spaced apart from each other to define at least one internal flow channel for channeling cooling air therein. A plenum is defined by an inner surface of the first sidewall and by an internal wall adjoining the flow channel. A pair of injection holes extend through the internal wall to feed the cooling air into the plenum. A plurality of film cooling holes extend through the first sidewall in flow communication with the plenum for channeling the cooling air in a film along an outer surface of the first sidewall to effect film cooling. The injection holes are aligned to converge toward each other so that cooling air jets discharged therefrom collide with each other inside the plenum to reduce pressure thereof. This regulates the pressure across the film cooling holes to improve film cooling.

Abstract translation: 燃气涡轮发动机翼型件包括在相对的前缘和后缘处连接在一起并且彼此间隔开的第一和第二侧壁，以限定至少一个用于在其中引导冷却空气的内部流动通道。 气室由第一侧壁的内表面和毗邻流动通道的内壁限定。 一对注射孔延伸穿过内壁，以将冷却空气送入气室。 多个膜冷却孔延伸穿过第一侧壁，与气室成流动连通，用于沿着第一侧壁的外表面引导薄膜中的冷却空气，以实现薄膜冷却。 喷射孔对准，朝向彼此会聚，使得排出的冷却空气射流彼此在气室内相互碰撞以减小压力。 这调节膜冷却孔的压力以改善膜冷却。

公开(公告)号：US5370499A

公开(公告)日：1994-12-06

申请号：US830144

申请日：1992-02-03

Applicant: Ching-Pang Lee

Inventor： Ching-Pang Lee

IPC: F01D5/18 ,  B63H1/14

CPC classification number: F01D5/187 ,  F01D5/186 ,  F05D2260/202 ,  F05D2260/2214 ,  Y02T50/676

Abstract: A turbine airfoil has a mesh cooling hole arrangement which includes first and second pluralities of cooling holes extending between internal and external surfaces of an airfoil side wall at least at a pressure side and extending from an internal chamber to the airfoil exterior. The cooling holes of each plurality extend generally parallel to one another. The cooling holes of the first and second pluralities intersect so as to define a plurality of spaced apart internal solid nodes in the side wall having pairs of opposite sides interconnected by pairs of opposite corners. The spaced nodes define a multiplicity of hole portions of the cooling holes which extend between and along opposite sides of adjacent nodes and a plurality of flow intersections which interconnect the hole portions of the cooling holes and are disposed between the corners of adjacent nodes. The sides of the nodes have lengths which are greater than the widths of the hole portions between adjacent nodes such that, when cooling fluid is passed through the cooling holes, jet flow actions are created through the hole portions which in turn generate jet interactions at the flow intersections to cause restriction of air flow and produce a pressure drop. Also, the cooling holes have flow inlets at the internal surface and flow outlets at the external surface of the airfoil side wall. The area of the flow inlets is substantially less than the area of the flow outlets.

Abstract translation: 涡轮机翼片具有网状冷却孔装置，其包括第一和第二多个冷却孔，其至少在压力侧在翼型侧壁的内表面和外表面之间延伸并且从内室延伸到翼型件外部。 每个多个冷却孔通常彼此平行地延伸。 第一和第二多个的冷却孔相交，以在侧壁中限定多个间隔开的内部实心节点，其具有通过成对的相对角互相连接的成对的相对侧。 间隔开的节点限定了冷却孔的多个孔部分，这些孔部分在相邻节点的相对侧面之间延伸并且沿相邻节点的相对侧延伸，以及多个流动交叉点，其互连冷却孔的孔部分并且设置在相邻节点的角部之间。 节点的侧面具有大于相邻节点之间的孔部分的宽度的长度，使得当冷却流体通过冷却孔时，通过孔部分产生喷射流动作用，这些孔部分又在其中产生射流相互作用 流动交叉点，以引起空气流动的限制并产生压降。 此外，冷却孔在内表面具有流入口，在翼型件侧壁的外表面具有流出口。 流入口的面积大大小于流出口的面积。

公开(公告)号：US5134844A

公开(公告)日：1992-08-04

申请号：US560538

申请日：1990-07-30

Applicant: Ching-Pang Lee ,  Clay K. Carlson ,  Monty L. Shelton ,  Harold P. Rieck, Jr. ,  Harvey W. Mason ,  Ambrose A. Hauser

Inventor： Ching-Pang Lee ,  Clay K. Carlson ,  Monty L. Shelton ,  Harold P. Rieck, Jr. ,  Harvey W. Mason ,  Ambrose A. Hauser

IPC: F02C7/00 ,  F01D5/08 ,  F02C7/18 ,  F02K3/06

CPC classification number: F01D5/08 ,  Y02T50/673 ,  Y02T50/676

Abstract: An aft entry system in a gas turbine engine is provided for cooling selected one or more of aft high and low pressure turbine stages of the engine. The aft entry system includes an air flow circuit for routing low pressure cooling air radially outward from an interstage region of a multi-stage compressor of a core engine of the gas turbine engine and axially in an aft direction to a plurality of radial passages, such as provided by stationary stator and outlet guide vanes, which provide an aft entry region to the selected high and low pressure turbine stages. The pressure of the cooling air routed by the air flow circuit of the aft entry system can be lower than the pressure of air at a discharge end of the compressor but higher than the pressure of the gas stream discharging from the selected turbine stage.

公开(公告)号：US20180318919A1

公开(公告)日：2018-11-08

申请号：US16036611

申请日：2018-07-16

Applicant: Ching-Pang Lee ,  Nan Jiang

Inventor： Ching-Pang Lee ,  Nan Jiang

IPC: B22D25/02 ,  F01D5/18 ,  B22C9/04 ,  B22C9/10 ,  B22C7/02 ,  B22D29/00 ,  B22C21/14 ,  B22C9/24

CPC classification number: B22D25/02 ,  B22C7/02 ,  B22C9/04 ,  B22C9/043 ,  B22C9/10 ,  B22C9/108 ,  B22C9/24 ,  B22C21/14 ,  B22D29/001 ,  F01D5/186 ,  F05D2220/32 ,  F05D2230/211

Abstract: A method of forming an airfoil (12), including: abutting end faces (72) of cantilevered film hole protrusions (64) extending from a ceramic core (50) against an inner surface (80) of a wax die (68) to hold the ceramic core in a fixed positional relationship with the wax die; casting an airfoil including a superalloy around the ceramic core; and machining film cooling holes (34) in the airfoil after the casting step to form an pattern of film cooling holes comprising the film cooling holes formed by the machining step and the cast film cooling holes (102) formed by the film hole protrusions during the casting step.

公开(公告)号：US09091495B2

公开(公告)日：2015-07-28

申请号：US13893392

申请日：2013-05-14

Applicant: Ching-Pang Lee

Inventor： Ching-Pang Lee

IPC: F28F13/12 ,  F01D5/18 ,  F04D29/38

CPC classification number: F28F13/12 ,  F01D5/187 ,  F05D2260/22141

Abstract: A cooling passage defined between first and second spaced apart sidewalls of a turbine engine component includes a turbulator system including a plurality of rows of turbulator members. Each row includes a first side turbulator member extending from the first sidewall, and a second side turbulator member extending from the second sidewall. The first and second side turbulator members are arranged such that a space is defined therebetween. The first and second side turbulator members are staggered with respect to one another such that respective forward and aft ends thereof are offset from one another. Each row further includes at least one elongate intermediate turbulator member located at least partially in the space between the respective first and second side turbulator members.

Abstract translation: 限定在涡轮发动机部件的第一和第二间隔开的侧壁之间的冷却通道包括包括多排湍流器构件的湍流器系统。 每排包括从第一侧壁延伸的第一侧湍流器构件和从第二侧壁延伸的第二侧湍流构件。 第一和第二侧面紊流器构件被布置成使得在它们之间限定空间。 第一和第二侧面紊流器构件相对于彼此交错，使得其相应的前端和后端彼此偏移。 每列还包括至少一个至少部分地位于相应的第一和第二侧面湍流器构件之间的空间中的细长的中间湍流器构件。

公开(公告)号：US09085981B2

公开(公告)日：2015-07-21

申请号：US13655675

申请日：2012-10-19

Applicant: Ching-Pang Lee ,  Jay A. Morrison

Inventor： Ching-Pang Lee ,  Jay A. Morrison

IPC: F01D9/06 ,  F23R3/02 ,  F23R3/06 ,  F23R3/54 ,  F01D9/02 ,  F23R3/00

CPC classification number: F01D9/023 ,  F05D2260/20 ,  F05D2260/202 ,  F23R3/002 ,  F23R3/06 ,  F23R3/54 ,  F23R2900/03045

Abstract: A ducting arrangement (10) for a can annular gas turbine engine, including: a duct (12, 14) disposed between a combustor (16) and a first row of turbine blades and defining a hot gas path (30) therein, the duct (12, 14) having raised geometric features (54) incorporated into an outer surface (80); and a flow sleeve (72) defining a cooling flow path (84) between an inner surface (78) of the flow sleeve (72) and the duct outer surface (80). After a cooling fluid (86) traverses a relatively upstream raised geometric feature (90), the inner surface (78) of the flow sleeve (72) is effective to direct the cooling fluid (86) toward a landing (94) separating the relatively upstream raised geometric feature (90) from a relatively downstream raised geometric feature (94).

Abstract translation: 一种用于罐式环形燃气涡轮发动机的管道装置（10），包括：设置在燃烧器（16）和第一排涡轮叶片之间并在其中限定热气体通道（30）的管道（12,14），所述管道 （12,14），其具有结合到外表面（80）中的凸起的几何特征（54）; 以及流动套筒（72），其限定在所述流动套筒（72）的内表面（78）和所述管道外表面（80）之间的冷却流动路径（84）。 在冷却流体（86）穿过相对上游的凸起的几何特征（90）之后，流动套筒（72）的内表面（78）有效地将冷却流体（86）引导到将相对 上游凸起几何特征（90）从相对下游凸起的几何特征（94）。

公开(公告)号：US09039357B2

公开(公告)日：2015-05-26

申请号：US14043958

申请日：2013-10-02

Applicant: Ching-Pang Lee

Inventor： Ching-Pang Lee

IPC: F04D29/08 ,  F01D11/00 ,  F01D11/02

CPC classification number: F01D11/001 ,  F01D11/02 ,  F05D2240/80 ,  F05D2250/37 ,  F05D2250/38 ,  F05D2250/71 ,  F05D2260/14

Abstract: A seal assembly between a disc cavity and a turbine section hot gas path includes a stationary vane assembly and a rotating blade assembly downstream from the vane assembly and including a plurality of blades that are supported on a platform and rotate with a turbine rotor and the platform during operation of the engine. The platform includes a radially outwardly facing first surface, a radially inwardly facing second surface, a third surface, and a plurality of grooves extending into the third surface. The grooves are arranged such that a space is defined between adjacent grooves. During operation of the engine, the grooves guide purge air out of the disc cavity toward the hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Abstract translation: 在盘腔和涡轮部分热气路径之间的密封组件包括固定叶片组件和在叶片组件下游的旋转叶片组件，并且包括支撑在平台上并与涡轮转子和平台一起旋转的多个叶片 在发动机运行期间。 平台包括径向向外的第一表面，径向向内的第二表面，第三表面和延伸到第三表面中的多个凹槽。 凹槽被布置成使得在相邻凹槽之间限定空间。 在发动机操作期间，凹槽将吹扫空气从盘腔朝向热气路径引导，使得净化空气相对于热气流通过热气路径的方向在所需方向上流动。

公开(公告)号：US20150122450A1

公开(公告)日：2015-05-07

申请号：US13998541

申请日：2013-11-07

Applicant: Ching-Pang Lee ,  Gerald L. Hillier ,  Jae Y. UM ,  Gm Salam Azad

Inventor： Ching-Pang Lee ,  Gerald L. Hillier ,  Jae Y. UM ,  Gm Salam Azad

IPC: B22C9/24 ,  B22C9/10

CPC classification number: B22C9/24 ,  B22C7/06 ,  B22C9/04 ,  B22C9/10 ,  F01D5/188 ,  F01D9/041 ,  F05D2260/201

Abstract: A cast ceramic core (110), including: an airfoil portion (116) shaped to define an inner surface (56) of an airfoil (52) of a vane segment (50); and a shell portion (122) having a backside-shaping surface (120) shaped to define a backside surface (68) of a shroud (62) of the vane segment. The backside-shaping surface has a higher elevation (132) and a lower elevation (134). The higher elevation is set apart from a nearest point (138) on the airfoil portion by the lower elevation. The airfoil portion and the shell portion are cast as a monolithic body during a single casting pour.

Abstract translation: 一种铸造陶瓷芯（110），包括：形成为限定叶片段（50）的翼型件（52）的内表面（56）的翼型部分（116） 以及壳体部分（122），其具有形成为限定叶片段的护罩（62）的后侧表面（68）的后侧成形表面（120）。 后侧成形表面具有较高的仰角（132）和较低的高度（134）。 较高的高度与翼面部分的最近点（138）分开，较低的高度。 翼型部分和壳体部分在单次铸造过程中作为整体式铸造。

公开(公告)号：US09017027B2

公开(公告)日：2015-04-28

申请号：US13760107

申请日：2013-02-06

Applicant: Christian X. Campbell ,  Ching-Pang Lee

Inventor： Christian X. Campbell ,  Ching-Pang Lee

IPC: F01D5/18 ,  F01D25/12 ,  F28F3/04 ,  F01D5/14 ,  F28F7/02

CPC classification number: F01D5/183 ,  F01D5/147 ,  F01D5/187 ,  F01D25/12 ,  F05D2240/304 ,  F05D2250/13 ,  F05D2260/2214 ,  F28F3/048 ,  F28F7/02

Abstract: A cooling channel (36, 36B, 63-66) cools inner surfaces (48, 50) of exterior walls (41, 43) of a component (20, 60). Interior side surfaces (52, 54) of the channel converge to a waist (W2), forming an hourglass shaped transverse profile (46). The inner surfaces (48, 50) may have fins (44) aligned with the coolant flow (22). The fins may have a transverse profile (56A, 56B) highest at mid-width of the inner surfaces (48, 50). Turbulators (92) may be provided on the side surfaces (52, 54) of the channel, and may urge the coolant flow toward the inner surfaces (48, 50). Each turbulator (92) may have a peak (97) that defines the waist of the cooling channel. Each turbulator may have a convex upstream side (93). These elements increase coolant flow in the corners (C) of the channel to more uniformly and efficiently cool the exterior walls (41, 43).

Abstract translation: 冷却通道（36,36B，63-66）冷却部件（20,60）的外壁（41,43）的内表面（48,50）。 通道的内侧表面（52,54）会聚到腰部（W2），形成沙漏形的横向轮廓（46）。 内表面（48,50）可具有与冷却剂流（22）对准的翅片（44）。 翅片可以具有在内表面（48,50）的中间宽度处最高的横向轮廓（56A，56B）。 紊流器（92）可以设置在通道的侧表面（52,54）上，并且可以促使冷却剂流向内表面（48,50）。 每个湍流器（92）可以具有限定冷却通道的腰部的峰（97）。 每个湍流器可以具有凸起的上游侧（93）。 这些元件增加通道的角部（C）中的冷却剂流动，以更均匀且有效地冷却外壁（41,43）。

公开(公告)号：US09017014B2

公开(公告)日：2015-04-28

申请号：US13930482

申请日：2013-06-28

Applicant: Ching-Pang Lee ,  Kok-Mun Tham ,  Eric Schroeder ,  Jamie Meeroff ,  Samuel R. Miller, Jr. ,  John J. Marra ,  Christian X. Campbell

Inventor： Ching-Pang Lee ,  Kok-Mun Tham ,  Eric Schroeder ,  Jamie Meeroff ,  Samuel R. Miller, Jr. ,  John J. Marra ,  Christian X. Campbell

IPC: F01D11/08 ,  F01D25/12 ,  F01D11/02 ,  F01D5/18

CPC classification number: F01D11/02 ,  F01D5/082 ,  F01D5/18 ,  F01D5/187 ,  F01D11/04 ,  F05D2260/14

Abstract: An outer rim seal arrangement (10), including: an annular rim (70) centered about a longitudinal axis (30) of a rotor disc (31), extending fore and having a fore-end (72), an outward-facing surface (74), and an inward-facing surface (76); a lower angel wing (62) extending aft from a base of a turbine blade (22) and having an aft end (64) disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (80); an upper angel wing (66) extending aft from the turbine blade base and having an aft end (68) disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (80, 82); and guide vanes (100) disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins (102) may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.

Abstract translation: 外缘密封装置（10），包括：以转子盘（31）的纵向轴线（30）为中心的环形边缘（70），其在前面延伸并具有前端（72），面向外的表面 （74）和面向内的表面（76）; 从天线叶片（22）的底部向后延伸的下部天使翼（62），并且具有设置在所述边缘向内相对表面的径向内侧的后端（64），以限定下部天使翼密封间隙（80）; 从所述涡轮叶片基部向后延伸的上天使翼（66），并且具有设置在所述轮缘朝外的表面的径向外侧的后端（68），以限定上天使翼密封间隙（80,82）; 以及设置在下天使翼密封间隙中的边缘向内表面上的导向叶片（100）。 抽吸翼（102）可以设置在上天使翼密封间隙中的上天使翼密封后端。