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71. 发明申请

US20080131265A1 Downstream plasma shielded film cooling 失效
标题翻译：下游等离子屏蔽膜冷却
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。

72. 发明申请

US20080131264A1 METHODS AND SYSTEM FOR COOLING INTEGRAL TURBINE SHROUD ASSEMBLIES 有权
标题翻译：用于冷却整体涡轮机组件的方法和系统
公开(公告)号：US20080131264A1
公开(公告)日：2008-06-05
申请号：US11565387
申请日：2006-11-30
申请人： Ching-Pang Lee , James Harvey Laflen , Dustrin Alfred Placke , George Elliott Moore , Katherine Jaynetorrence Andersen , Daniel Verner Jones
发明人： Ching-Pang Lee , James Harvey Laflen , Dustrin Alfred Placke , George Elliott Moore , Katherine Jaynetorrence Andersen , Daniel Verner Jones
IPC分类号： F01D25/14 , F01D25/28
CPC分类号： F01D11/24 , F01D25/246 , F05D2260/201 , F05D2260/205 , Y02T50/671 , Y02T50/676
摘要： A method for cooling a turbine shroud assembly includes providing a turbine shroud assembly including a shroud segment having a leading edge, a trailing edge and a midsection defined therebetween. A shroud support circumferentially spans and supports the shroud segment. The shroud support includes a forward hanger coupled to the leading edge, a midsection hanger coupled to the midsection and an aft hanger coupled to the trailing edge. An annular shroud ring structure includes a midsection position control ring coupled to the midsection hanger and an aft position control ring coupled to the aft hanger. Cooling air is extracted from a compressor positioned upstream of the turbine shroud assembly. Cooling air is metered through the shroud support directly into only at least one active convection cooling zone defined between the shroud segment and the shroud support while substantially preventing cooling air from entering an inactive convection cooling zone positioned radially outwardly from the at least one active convection cooling zone and defined between the shroud support and the shroud ring structure and between the midsection position control ring and the aft position control ring.
摘要翻译：一种用于冷却涡轮机护罩组件的方法包括提供一种涡轮机护罩组件，其包括具有前缘，后缘和在其间限定的中部的护罩段。护罩支撑件周向地跨越并支撑护罩段。护罩支撑件包括联接到前缘的前悬挂器，联接到中部的中间吊架和耦合到后缘的后挂架。环形护罩环结构包括连接到中间悬挂架的中央位置控制环和联接到后悬挂架的后部位置控制环。从位于涡轮机罩组件上游的压缩机抽出冷却空气。冷却空气通过护罩支撑件计量直接仅限于限定在护罩区段和护罩支撑件之间的至少一个主动对流冷却区域，同时基本上防止冷却空气进入从至少一个主动对流冷却器径向向外定位的非活动对流冷却区域并且限定在护罩支撑件和护罩环结构之间以及中间位置控制环和后部位置控制环之间。

73. 发明申请

US20080131260A1 METHOD AND SYSTEM TO FACILITATE COOLING TURBINE ENGINES 有权
标题翻译：冷却涡轮发动机的方法和系统
公开(公告)号：US20080131260A1
公开(公告)日：2008-06-05
申请号：US11565229
申请日：2006-11-30
申请人： Ching-Pang Lee , Eric Alan Estill , James Harvey Laflen , Paul Hadley Vitt , Michael Elliot Wymore
发明人： Ching-Pang Lee , Eric Alan Estill , James Harvey Laflen , Paul Hadley Vitt , Michael Elliot Wymore
IPC分类号： F01D25/12
CPC分类号： F01D25/12 , F01D9/04 , F01D11/08 , F01D11/24 , F01D25/246 , F05D2240/11 , F05D2240/81 , F05D2250/314 , Y02T50/671 , Y02T50/676 , Y10T29/49236
摘要： A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.
摘要翻译：提供一种组装燃气涡轮发动机的方法。该方法包括联接燃气涡轮发动机内的至少一个涡轮喷嘴段。至少一个涡轮喷嘴段包括在内带和外带之间延伸的至少一个翼型叶片，其包括后翼缘和径向内表面。该方法还包括将至少一个涡轮机护罩段联接到至少一个涡轮喷嘴段的下游，其中至少一个涡轮机护罩段包括前缘和径向内表面，并且将冷却流体源与至少一个涡轮喷嘴段，使得被引导到每个涡轮喷嘴外带后缘的冷却流体朝向至少一个涡轮机护罩段的前缘以倾斜的排出角度引导。

74. 发明申请

US20080080968A1 METHODS AND APPARATUS FOR ASSEMBLING TURBINE ENGINES 有权
标题翻译：组装涡轮发动机的方法和装置
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。

75. 发明申请

US20070134099A1 Damper cooled turbine blade 有权
标题翻译：阻尼器冷却涡轮叶片
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。

76. 发明申请

US20070134089A1 Methods and apparatus for assembling turbine engines 有权
标题翻译：涡轮发动机组装方法及装置
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。

77. 发明申请

US20070104571A1 Methods and apparatus for assembling turbine engines 失效
标题翻译：涡轮发动机组装方法及装置
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。

78. 发明申请

US20070031245A1 Thermally compliant C-clip 有权
标题翻译：耐热兼容的C形夹
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。

79. 发明授权

US07118342B2 Fluted tip turbine blade 有权
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.

80. 发明授权

US07094444B2 Method for repairing coated components using NiAl bond coats 有权
标题翻译：使用NiAl粘合涂层修复涂层部件的方法
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。

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