公开(公告)号：US20070104571A1

公开(公告)日：2007-05-10

申请号：US11271101

申请日：2005-11-10

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

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

IPC: F01D9/00

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

Abstract: 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.

Abstract translation: 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

Applicant: Ching-Pang Lee ,  Chander Prakash

Inventor： Ching-Pang Lee ,  Chander Prakash

IPC: F01D5/18

CPC classification number: F01D5/20 ,  F05D2250/712 ,  Y02T50/67

Abstract: 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.

Abstract translation: 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

Applicant: Michael Ruthemeyer ,  Glenn Nichols ,  Ching-Pang Lee

Inventor： Michael Ruthemeyer ,  Glenn Nichols ,  Ching-Pang Lee

IPC: F01D25/26

CPC classification number: F01D9/04 ,  F05D2230/60 ,  F05D2240/11 ,  F05D2260/30 ,  F05D2260/941

Abstract: 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.

Abstract translation: 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

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

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

IPC: F01D5/14

CPC classification number: F01D5/186 ,  F01D5/20 ,  F01D11/08 ,  Y02T50/676

Abstract: 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

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

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

IPC: B05C13/00 ,  B05D1/36

CPC classification number: 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

Abstract: 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.

Abstract translation: 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。

公开(公告)号：US20060140768A1

公开(公告)日：2006-06-29

申请号：US11022121

申请日：2004-12-24

Applicant: Anna Tam ,  Ching-Pang Lee ,  Kevin Kirtley ,  Ronald Bunker ,  Scott Lamson ,  Scott Carson

Inventor： Anna Tam ,  Ching-Pang Lee ,  Kevin Kirtley ,  Ronald Bunker ,  Scott Lamson ,  Scott Carson

IPC: F01D11/00

CPC classification number: F01D5/145 ,  F01D5/143 ,  F05D2240/80 ,  F05D2250/70 ,  F05D2250/711 ,  F05D2250/712 ,  Y02T50/673

Abstract: A turbine stage includes a row of airfoils joined to corresponding platforms to define flow passages therebetween. Each airfoil includes opposite pressure and suction sides and extends in chord between opposite leading and trailing edges. Each platform has a scalloped flow surface including a bulge adjoining the pressure side adjacent the leading edge, and a bowl adjoining the suction side aft of the leading edge.

公开(公告)号：US20060053800A1

公开(公告)日：2006-03-16

申请号：US10941546

申请日：2004-09-15

Applicant: Robert Orlando ,  Kattalaicheri Venkataramani ,  Ching-Pang Lee ,  Thomas Moniz ,  Kurt Murrow

Inventor： Robert Orlando ,  Kattalaicheri Venkataramani ,  Ching-Pang Lee ,  Thomas Moniz ,  Kurt Murrow

IPC: F02C7/12

CPC classification number: F02C7/18 ,  F02C5/00 ,  F05D2260/20 ,  F05D2260/96 ,  F23R3/005 ,  F23R3/04 ,  F23R7/00 ,  F23R2900/03044 ,  F23R2900/03281 ,  Y02T50/676

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the compressor, where the core system further includes an intermediate compressor positioned downstream of and in flow communication with the booster compressor, the intermediate compressor being connected to a second drive shaft, and a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, a low pressure turbine positioned downstream of and in flow communication with the core system, the low pressure turbine being utilized to power the first drive shaft. The core system may also include an intermediate turbine positioned downstream of the combustion system in flow communication with the working fluid, where the intermediate turbine is utilized to power the second drive shaft. A first source of compressed air having a predetermined pressure is provided to the combustion system inlet and a second source of compressed air having a pressure greater than the first source of compressed air is provided to cool the combustion system.

Abstract translation: 一种燃气涡轮发动机，其具有穿过其的纵向中心线轴线，包括：在所述燃气涡轮发动机的前端处的风扇部分，至少包括连接到第一驱动轴的第一风扇叶片排; 位于与包括多个级的风扇部分的下游并且至少部分流动连通的增压压缩机，每个级包括固定压缩机叶片排和连接到驱动轴的旋转压缩机叶片排，并与固定压缩机叶片排相互指向 ; 位于压缩机下游的核心系统，其中所述核心系统还包括位于所述增压压缩机下游并与所述增压压缩机流动连通的中间压缩机，所述中间压缩机连接到第二驱动轴，以及用于产生气体脉冲的燃烧系统 从提供给其入口的流体流增加压力和温度，以便在出口处产生工作流体; 以及低压涡轮机，其位于所述核心系统的下游并与所述核心系统流动连通，所述低压涡轮机用于为所述第一驱动轴提供动力。 核心系统还可以包括位于燃烧系统下游的中间涡轮机，其与工作流体流动连通，其中中间涡轮机用于为第二驱动轴提供动力。 具有预定压力的第一压缩空气源被提供给燃烧系统入口，并且提供具有大于第一压缩空气源的压力的第二压缩空气源以冷却燃烧系统。

公开(公告)号：US20060051209A1

公开(公告)日：2006-03-09

申请号：US10937642

申请日：2004-09-09

Applicant: Ching-Pang Lee ,  James Bailey ,  Nicholas Palmer ,  Wenfeng Lu

Inventor： Ching-Pang Lee ,  James Bailey ,  Nicholas Palmer ,  Wenfeng Lu

IPC: F01D5/18

CPC classification number: F01D5/186 ,  F01D5/20 ,  F01D11/08 ,  Y02T50/676

Abstract: 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.

Abstract translation: 涡轮机叶片包括具有在前缘和后缘之间延伸的压力和吸力侧壁以及根部和尖端的翼型件。 尖端包括从形成开口尖端腔的尖端底部延伸的尖叫肋。 沿着压力侧壁的肋具有平方的外角，并且凹槽沿尖端底部的肋的基部延伸。

公开(公告)号：US20060029723A1

公开(公告)日：2006-02-09

申请号：US10714430

申请日：2003-11-13

Applicant: Joseph Rigney ,  Ching-Pang Lee ,  Ramgopal Darolia

Inventor： Joseph Rigney ,  Ching-Pang Lee ,  Ramgopal Darolia

IPC: C23C16/52 ,  B05D3/00

CPC classification number: 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

Abstract: 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.

Abstract translation: 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。

公开(公告)号：US06990797B2

公开(公告)日：2006-01-31

申请号：US10656518

申请日：2003-09-05

Applicant: Kattalaicheri Srinivasan Venkataramani ,  Lawrence Butler ,  Ching-Pang Lee

Inventor： Kattalaicheri Srinivasan Venkataramani ,  Lawrence Butler ,  Ching-Pang Lee

IPC: F02C7/0472 ,  F02K3/04

CPC classification number: F02C7/047 ,  F05D2260/208 ,  Y02T50/672 ,  Y02T50/675

Abstract: A method facilitates assembling a turbine engine to facilitate preventing ice accumulation on the turbine engine during engine operation. The method comprises coupling at least one heat pipe to the engine such that a first end of the at least one heat pipe is coupled in thermal communication with a heat source, and coupling a second end of the at least one heat pipe in thermal communication with an outer surface of an engine component that is upstream from the heat source.

Abstract translation: 一种方法有利于组装涡轮发动机以便于在发动机运转期间防止在涡轮发动机上的积冰。 该方法包括将至少一个热管耦合到发动机，使得至少一个热管的第一端与热源热连通地联接，并将至少一个热管的第二端与热源连通 发动机部件的外表面，其位于热源的上游。