公开(公告)号：US20080135202A1

公开(公告)日：2008-06-12

申请号：US11567477

申请日：2006-12-06

申请人： Ching-Pang Lee ,  Hsin-Pang Wang ,  Ram Kumar Upadhyay ,  Paul Richard Myers ,  Marc Thomas Edgar ,  Thomas Donald Martyn ,  Eric Alan Estill

发明人： Ching-Pang Lee ,  Hsin-Pang Wang ,  Ram Kumar Upadhyay ,  Paul Richard Myers ,  Marc Thomas Edgar ,  Thomas Donald Martyn ,  Eric Alan Estill

IPC分类号： B22C9/10 ,  B28B7/34

CPC分类号： B22C9/10 ,  B22C9/101 ,  B22C9/103 ,  B28B7/342 ,  B28B7/346

摘要： Disclosed herein is a composite core die comprising a reusable core die; and a disposable core die; wherein the disposable core die is in physical communication with the reusable core die; and further wherein surfaces of communication between the disposable core die and the reusable core die serve as barriers to prevent the leakage of a slurry that is disposed in the composite core die.

摘要翻译： 本文公开了一种复合芯模，其包括可重复使用的芯模; 和一次性芯模; 其中所述一次性芯管芯与所述可重复使用的芯管芯物理连通; 并且其中一次性芯模和可重复使用的芯模之间的连通表面用作阻挡物以防止布置在复合芯模中的浆料的泄漏。

公开(公告)号：US20080131265A1

公开(公告)日：2008-06-05

申请号：US11606853

申请日：2006-11-30

申请人： Ching-Pang Lee ,  Aspi Rustom Wadia ,  David Glenn Cherry ,  Je-Chin Han

发明人： Ching-Pang Lee ,  Aspi Rustom Wadia ,  David Glenn Cherry ,  Je-Chin Han

IPC分类号： F01D5/18 ,  F01D25/12

CPC分类号： F01D5/186 ,  F05D2260/202 ,  F05D2260/221 ,  F05D2270/172 ,  F15D1/12 ,  F23R3/002 ,  F23R2900/03042 ,  Y02T50/673 ,  Y02T50/676 ,  Y10S165/903 ,  Y10S415/914

摘要： An downstream plasma boundary layer shielding system includes film cooling apertures disposed through a wall having cold and hot surfaces and angled in a downstream direction from a cold surface of the wall to an outer hot surface of the wall. A plasma generator located downstream of the film cooling apertures is used for producing a plasma extending downstream over the film cooling apertures. Each plasma generator includes inner and outer electrodes separated by a dielectric material disposed within a groove in the outer hot surface. The wall may be part of a hollow airfoil or an annular combustor or exhaust liner. A method for operating the downstream plasma boundary layer shielding system includes forming a plasma extending in the downstream direction over the film cooling apertures along the outer hot surface of the wall. The method may further include operating the plasma generator in steady state or unsteady modes.

摘要翻译： 下游等离子体边界层屏蔽系统包括薄膜冷却孔，该薄膜冷却孔通过具有冷和热表面的壁布置，并且在下游方向上从壁的冷表面到壁的外部热表面成角度。 位于膜冷却孔下游的等离子体发生器用于产生在膜冷却孔的下游延伸的等离子体。 每个等离子体发生器包括由设置在外部热表面中的凹槽内的电介质材料分开的内部和外部电极。 该壁可以是中空翼型件或环形燃烧器或排气衬套的一部分。 一种用于操作下游等离子体边界层屏蔽系统的方法包括形成沿着下游方向延伸的等离子体沿着壁的外部热表面的膜冷却孔。 该方法还可以包括以等稳态或不稳定模式操作等离子体发生器。

公开(公告)号：US20080080968A1

公开(公告)日：2008-04-03

申请号：US11538273

申请日：2006-10-03

申请人： Joseph Michael Guentert ,  Ching-Pang Lee

发明人： Joseph Michael Guentert ,  Ching-Pang Lee

IPC分类号： F01D25/26

CPC分类号： F01D25/246 ,  F01D11/005 ,  F05D2240/57

摘要： A method for assembling a gas turbine engine is provided. The method includes providing a turbine nozzle including an outer band and an inner band, wherein each band includes a leading edge, a trailing edge, and a body extending therebetween. At least one of the outer band and the inner band has at least one radial tab extending outward therefrom. The method also includes coupling at least one seal between at least one of the radial tabs extending from the outer band and the inner band and a respective leading edge of the outer and inner band. The method also includes positioning at least one non-planar seal support against at least one portion of the seal.

摘要翻译： 提供一种用于组装燃气涡轮发动机的方法。 该方法包括提供包括外带和内带的涡轮喷嘴，其中每个带包括前缘，后缘和在其间延伸的本体。 外带和内带中的至少一个具有从其向外延伸的至少一个径向突片。 该方法还包括在从外带和内带延伸的至少一个径向突片和外带和内带的相应前缘之间联接至少一个密封件。 该方法还包括将至少一个非平面密封支撑件定位在密封件的至少一部分上。

公开(公告)号：US20070134099A1

公开(公告)日：2007-06-14

申请号：US11296910

申请日：2005-12-08

申请人： Ching-Pang Lee ,  Steven Brassfield ,  Jan Schilling

发明人： Ching-Pang Lee ,  Steven Brassfield ,  Jan Schilling

IPC分类号： F01D11/00

CPC分类号： F01D5/22 ,  F01D11/006 ,  F05D2240/81 ,  F05D2250/292 ,  F05D2250/314 ,  Y02T50/671 ,  Y02T50/676 ,  Y10S416/50

摘要： A turbine blade includes an airfoil, platform, shank, and dovetail integrally joined together. A cooling chamber is located under the platform and has a portal exposed outwardly from the shank. A damper seat surrounds the portal and is recessed under the platform for receiving a vibration damper to sealingly close the chamber across the portal.

摘要翻译： 涡轮机叶片包括整体连接在一起的翼型件，平台，柄和燕尾榫。 冷却室位于平台下方并具有从柄部向外露出的入口。 阻尼器座圈围绕入口并且凹陷在平台下方，用于接收振动阻尼器以密封地封闭穿过入口的腔室。

公开(公告)号：US20070134089A1

公开(公告)日：2007-06-14

申请号：US11297699

申请日：2005-12-08

申请人： Ching-Pang Lee ,  Joseph Guentert ,  Wenfeng Lu ,  Mitchell Iles

发明人： Ching-Pang Lee ,  Joseph Guentert ,  Wenfeng Lu ,  Mitchell Iles

IPC分类号： F01D9/00

CPC分类号： F01D9/041 ,  F01D5/143 ,  F01D5/145 ,  F01D9/02 ,  F05D2240/81 ,  Y02T50/673 ,  Y02T50/676

摘要： A method facilitates the assembly of a gas turbine engine. The method of assembly comprises providing a turbine nozzle including an inner band, an outer band, at least one vane extending between the inner and outer bands, and at least one leading edge fillet extending between the at least one vane and at least one of the inner and outer bands, wherein a leading edge of the at least one vane is downstream from the leading edges of the inner and outer bands, and coupling the turbine nozzle within the gas turbine engine such that the leading edge fillet is configured to facilitate minimizing vortex formation along the vane leading edge adjacent at least one of the inner and outer bands.

摘要翻译： 一种方法有助于燃气涡轮发动机的组装。 组装方法包括提供涡轮喷嘴，该涡轮喷嘴包括内带，外带，在内带和外带之间延伸的至少一个叶片，以及在至少一个叶片和至少一个叶片之间延伸的至少一个前缘圆角 内部和外部带，其中所述至少一个叶片的前缘在所述内部和外部带的前缘的下游，以及将所述涡轮喷嘴连接在所述燃气涡轮发动机内，使得所述前缘圆角被配置为有助于最小化涡旋 沿着与内带和外带中的至少一个相邻的叶片前缘的形成。

公开(公告)号：US20070104571A1

公开(公告)日：2007-05-10

申请号：US11271101

申请日：2005-11-10

申请人： Vincent Drerup ,  Harold Hansel ,  James Ketzer ,  Ching-Pang Lee

发明人： Vincent Drerup ,  Harold Hansel ,  James Ketzer ,  Ching-Pang Lee

IPC分类号： F01D9/00

CPC分类号： F02C7/28 ,  F01D11/005 ,  F05D2240/12 ,  F05D2260/30 ,  Y10T29/49297 ,  Y10T29/4932 ,  Y10T29/49323

摘要： A method for assembling a gas turbine engine is provided. The method comprises coupling a first turbine nozzle within the engine, coupling a second turbine nozzle circumferentially adjacent the first turbine nozzle such that a gap is defined between the first and second turbine nozzles and providing at least one spline seal including a substantially planar body. The method also comprises forming at least one retainer tab to extend outward from the body portion of the at least one spline seal, and inserting the at least one spline seal into a slot defined in at least one of the first and second turbine nozzles to facilitate reducing leakage through said gap, such that the at least one retainer tab facilitates retaining the retainer tab within the turbine nozzle slot.

摘要翻译： 提供一种用于组装燃气涡轮发动机的方法。 该方法包括将第一涡轮喷嘴联接在发动机内，将第二涡轮喷嘴与第一涡轮喷嘴周向相邻地连接，使得在第一涡轮喷嘴和第二涡轮喷嘴之间限定间隙，并提供包括基本上平面的主体的至少一个花键密封。 该方法还包括形成至少一个保持器突片以从至少一个花键密封件的主体部分向外延伸，以及将至少一个花键密封件插入限定在第一和第二涡轮喷嘴中的至少一个中的狭槽中，以便于 减少通过所述间隙的泄漏，使得至少一个保持器突片有助于将保持器突片保持在涡轮喷嘴槽内。

公开(公告)号：US20070059173A1

公开(公告)日：2007-03-15

申请号：US11162434

申请日：2005-09-09

申请人： Ching-Pang Lee ,  Chander Prakash

发明人： Ching-Pang Lee ,  Chander Prakash

IPC分类号： F01D5/18

CPC分类号： F01D5/20 ,  F05D2250/712 ,  Y02T50/67

摘要： An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil. At least a portion of the pressure-side tip wall is recessed from the pressure sidewall to define an outwardly facing tip shelf, such that the pressure-side tip wall and the tip shelf define a trough therebetween.

摘要翻译： 用于燃气涡轮发动机的翼型件包括根部，尖端，前缘，后缘以及大致沿径向轴线延伸的相对的压力和吸力侧壁。 翼型件包括在压力侧和吸力侧壁之间延伸的顶盖; 并且间隔开的吸入侧和压力侧尖端壁从尖端帽径向向外延伸以在它们之间限定尖端腔。 压力侧末端壁包括连续凹入的弯曲弧形部分，其至少一部分从翼型件的径向轴向向外延伸。 压力侧顶端壁的至少一部分从压力侧壁凹入以限定向外的顶端架，使得压力侧顶端壁和尖端架在其间限定槽。

公开(公告)号：US20070031245A1

公开(公告)日：2007-02-08

申请号：US11161518

申请日：2005-08-06

申请人： Michael Ruthemeyer ,  Glenn Nichols ,  Ching-Pang Lee

发明人： Michael Ruthemeyer ,  Glenn Nichols ,  Ching-Pang Lee

IPC分类号： F01D25/26

CPC分类号： F01D9/04 ,  F05D2230/60 ,  F05D2240/11 ,  F05D2260/30 ,  F05D2260/941

摘要： A C-clip for a gas turbine engine includes an arcuate outer arm having a first radius of curvature; an arcuate, inner arm having a second radius of curvature which is substantially greater than the first radius of curvature; and an arcuate extending flange connecting the outer and inner arms. The flange, the outer arm, and the inner arm collectively define a generally C-shaped cross-section. A shroud assembly includes a shroud segment with a mounting flange, and a shroud hanger with an arcuate hook disposed in mating relationship to the mounting flange. An arcuate C-clip having inner and outer arms overlaps the hook and the mounting flange. The shroud segment and the C-clip are subject to thermal expansion at the hot operating condition. A dimension of one of the shroud segment and the C-clip are selected to produce a preselected dimensional relationship therebetween at the hot operating condition.

摘要翻译： 用于燃气涡轮发动机的C形夹包括具有第一曲率半径的弓形外臂; 具有第二曲率半径的弓形内臂，其基本上大于所述第一曲率半径; 以及连接外臂和内臂的弓形延伸法兰。 凸缘，外臂和内臂共同地限定了大致C形的横截面。 护罩组件包括具有安装凸缘的护罩段和具有与安装法兰配合关系设置的弓形钩的护罩。 具有内臂和外臂的弧形C形夹与钩和安装法兰重叠。 护罩段和C形夹在热操作条件下经受热膨胀。 选择护罩段和C形夹之一的尺寸以在热操作条件下产生预定尺寸关系。

公开(公告)号：US07118342B2

公开(公告)日：2006-10-10

申请号：US10937642

申请日：2004-09-09

申请人： Ching-Pang Lee ,  James Robert Bailey ,  Nicholas Charles Palmer ,  Wenfeng Lu

发明人： Ching-Pang Lee ,  James Robert Bailey ,  Nicholas Charles Palmer ,  Wenfeng Lu

IPC分类号： F01D5/14

CPC分类号： F01D5/186 ,  F01D5/20 ,  F01D11/08 ,  Y02T50/676

摘要： A turbine blade includes an airfoil having pressure and suction sidewalls extending between leading and trailing edges and root and tip. The tip includes squealer ribs extending from a tip floor forming an open tip cavity. The rib along the pressure sidewall has a squared external corner, and a flute extends along the base of the rib at the tip floor.

公开(公告)号：US07094444B2

公开(公告)日：2006-08-22

申请号：US10714430

申请日：2003-11-13

申请人： Joseph D. Rigney ,  Ching-Pang Lee ,  Ramgopal Darolia

发明人： Joseph D. Rigney ,  Ching-Pang Lee ,  Ramgopal Darolia

IPC分类号： B05C13/00 ,  B05D1/36

CPC分类号： B23P6/002 ,  C23C10/02 ,  C23C10/60 ,  C23C28/321 ,  C23C28/3215 ,  C23C28/345 ,  C23C28/3455 ,  C23C28/36 ,  C25D3/50 ,  C25D7/00 ,  C25D7/008 ,  F05B2230/31 ,  F05B2230/90 ,  F05C2253/12

摘要： According to an embodiment of the invention, a method for repairing a coated high pressure turbine blade, which has been exposed to engine operation, to restore coated airfoil contour dimensions of the blade, and improve upon the prior bond coat is disclosed. The method comprises providing an engine run high pressure turbine blade including a base metal substrate made of a nickel-based alloy and having thereon a thermal barrier coating system. The thermal barrier coating system comprises a diffusion bond coat on the base metal substrate and a top ceramic thermal barrier coating comprising a yttria stabilized zirconia material. The top ceramic thermal barrier coating has a nominal thickness t. The method further comprises removing the thermal barrier coating system, wherein a portion of the base metal substrate also is removed, and determining the thickness of the base metal substrate removed. The portion of the base metal substrate removed has a thickness, Δt. The method also comprises applying a β phase NiAl overlay coating to the substrate, and determining the difference in thickness, Δx, between the β phase NiAl overlay coating and the previously removed bond coat. The method further comprises reapplying the top ceramic thermal barrier coating to a nominal thickness of t+Δt−Δx, wherein Δt compensates for the portion of removed base metal substrate. Advantageously, the coated airfoil contour dimensions of the high pressure turbine blade are restored to about the coated dimensions preceding the engine run.

摘要翻译： 根据本发明的实施例，公开了一种已经暴露于发动机操作的用于修复涂覆的高压涡轮机叶片以恢复叶片的被覆翼型轮廓尺寸并改进先前的粘结涂层的方法。 该方法包括提供一种发动机运行的高压涡轮叶片，其包括由镍基合金制成的基底金属基底并且具有热障涂层系统。 热障涂层系统包括在基底金属基底上的扩散粘合涂层和包含氧化钇稳定的氧化锆材料的顶部陶瓷热障涂层。 顶部陶瓷热障涂层具有标称厚度t。 该方法还包括去除热障涂层系统，其中基底金属衬底的一部分也被去除，并且确定移除的母体金属衬底的厚度。 去除的贱金属基材的部分厚度为Deltat。 该方法还包括将β相NiAl覆盖涂层施加到基底上，并确定β相NiAl覆盖涂层与先前去除的粘结涂层之间的厚度差异Deltax。 该方法还包括将顶部陶瓷热障涂层重新施加到t + Deltat-Deltax的标称厚度，其中Deltat补偿去除的基底金属基底的部分。 有利地，高压涡轮机叶片的被覆翼型轮廓尺寸恢复到发动机运行之前的涂层尺寸。