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1. 发明授权

US07186091B2 Methods and apparatus for cooling gas turbine engine components 失效
标题翻译：用于冷却燃气涡轮发动机部件的方法和装置
公开(公告)号：US07186091B2
公开(公告)日：2007-03-06
申请号：US10984292
申请日：2004-11-09
申请人： Ching-Pang Lee , Ronald Scott Bunker , Harvey Michael Maclin , Ramgopal Darolia
发明人： Ching-Pang Lee , Ronald Scott Bunker , Harvey Michael Maclin , Ramgopal Darolia
IPC分类号： F01D5/18
CPC分类号： F23R3/002 , F01D5/183 , F01D5/186 , F01D5/288 , F05D2230/90 , F05D2300/611 , F23R2900/03041
摘要： A method of cooling a gas turbine engine component having a perforate metal wall includes providing a plurality of pores in the wall, wherein the pores extend substantially perpendicularly through the wall, and wherein the pores are covered and sealed closed at first ends thereof by a thermal barrier coating disposed over a first surface of the wall, and providing a plurality of film cooling holes in the wall, wherein the holes extend substantially perpendicularly through the wall and the thermal barrier coating. The method also includes providing cooling fluid to the plurality of pores and the plurality of film cooling holes along a second surface of the wall, channeling the cooling fluid through the pores for back side cooling an inner surface of the thermal barrier coating, and channeling the cooling fluid through the holes for film cooling an outer surface of the thermal barrier coating.
摘要翻译：一种冷却具有穿孔金属壁的燃气涡轮发动机部件的方法包括在壁中提供多个孔，其中孔基本上垂直延伸穿过壁，并且其中孔在其第一端被热封屏障涂层，其设置在所述壁的第一表面上，并且在所述壁中提供多个膜冷却孔，其中所述孔基本上垂直延伸穿过所述壁和所述热障涂层。该方法还包括沿着壁的第二表面向多个孔和多个膜冷却孔提供冷却流体，将冷却流体引导通过孔，用于背面冷却热障涂层的内表面，并且引导冷却流体通过孔，用于薄膜冷却隔热涂层的外表面。

2. 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。

3. 发明申请

US20060029723A1 METHOD FOR REPAIRING COATED COMPONENTS USING NIAL BOND COATS 有权
标题翻译：使用NIAL BOND COATS修复涂层组件的方法
公开(公告)号：US20060029723A1
公开(公告)日：2006-02-09
申请号：US10714430
申请日：2003-11-13
申请人： Joseph Rigney , Ching-Pang Lee , Ramgopal Darolia
发明人： Joseph Rigney , Ching-Pang Lee , Ramgopal Darolia
IPC分类号： C23C16/52 , B05D3/00
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。

4. 发明申请

US20050271514A1 COATING AND COATING PROCESS INCORPORATING RAISED SURFACE FEATURES FOR AN AIR-COOLED SURFACE 有权
标题翻译：包含空气冷却表面的提高表面特性的涂层和涂层工艺
公开(公告)号：US20050271514A1
公开(公告)日：2005-12-08
申请号：US10906903
申请日：2005-03-11
申请人： Ching-Pang Lee , Robert Schafrik , Ramgopal Darolia , Joseph Rigney
发明人： Ching-Pang Lee , Robert Schafrik , Ramgopal Darolia , Joseph Rigney
IPC分类号： B32B15/16 , C08F2/48 , C23C4/02 , C23C4/08 , C23C14/02 , C23C14/04 , C23C14/16 , C23C16/06 , C23C28/00 , F23M5/00 , F23R3/00 , F28F13/18
CPC分类号： C23C4/02 , C23C4/073 , C23C4/08 , C23C14/025 , C23C14/044 , C23C14/16 , C23C28/021 , C23C28/022 , F05B2240/11 , F05B2250/60 , F05B2260/224 , F05C2253/12 , F23M5/00 , F23M2900/05004 , F23R3/002 , F28F13/18 , F28F13/185 , Y10T428/12063 , Y10T428/12493 , Y10T428/12639 , Y10T428/12646 , Y10T428/12931 , Y10T428/12944 , Y10T428/12993
摘要： A coating and coating process for incorporating surface features on an air-cooled substrate surface of a component for the purpose of promoting heat transfer from the component. The coating process generally comprises depositing a first metallic coating material on the surface of the component using a first set of coating conditions to form a first environmental coating layer, and then depositing a second metallic coating material using a second set of coating conditions that differ from the first set, such that an outer environmental coating layer is formed having raised surface features that cause the surface of the outer environmental coating layer to be rougher than the surface of the first environmental coating layer.
摘要翻译：用于将部件的空气冷却基板表面上的表面特征引入到涂层和涂覆方法中，以促进从部件的热传递。涂覆方法通常包括使用第一组涂覆条件在组件的表面上沉积第一金属涂层材料以形成第一环境涂层，然后使用第二组涂覆条件沉积第二金属涂层材料第一组，使得外部环境涂层形成具有使外部环境涂层的表面比第一环境涂层的表面更粗糙的凸起表面特征。

5. 发明申请

US20050106316A1 Method for repairing coated components 有权
标题翻译：修复涂层部件的方法
公开(公告)号：US20050106316A1
公开(公告)日：2005-05-19
申请号：US10714213
申请日：2003-11-13
申请人： Joseph Rigney , Ching-Pang Lee , Ramgopal Darolia
发明人： Joseph Rigney , Ching-Pang Lee , Ramgopal Darolia
IPC分类号： B23P6/00 , C23C26/00 , C23C28/00 , F01D5/00 , F01D5/14 , F01D5/28 , F01D25/00 , F02C7/00 , B32B35/00
CPC分类号： C23C28/321 , C23C26/00 , C23C28/325 , C23C28/3455 , F01D5/005 , F01D5/288 , F05D2230/80 , F05D2230/90 , F05D2300/501 , Y10T29/49318
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, 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 reapplying the diffusion bond coat to the substrate, wherein the bond coat is reapplied to a thickness, which is about the same as applied prior to the engine operation; and reapplying the top ceramic thermal barrier coating to a nominal thickness of t+Δt, 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。该方法还包括将扩散粘合涂层重新施加到基底上，其中粘结涂层重新施加到与发动机操作前相同的厚度; 并将顶部陶瓷热障涂层重新施加到标称厚度t + Deltat，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。

6. 发明授权

US06617003B1 Directly cooled thermal barrier coating system 失效
标题翻译：直接冷却热障涂层系统
公开(公告)号：US06617003B1
公开(公告)日：2003-09-09
申请号：US09707023
申请日：2000-11-06
申请人： Ching-Pang Lee , Robert Edward Schafrik , Ramgopal Darolia
发明人： Ching-Pang Lee , Robert Edward Schafrik , Ramgopal Darolia
IPC分类号： B32B1504
CPC分类号： C23C28/3215 , C23C28/321 , C23C28/325 , C23C28/3455 , F01D5/182 , F01D5/187 , F01D5/288 , Y02T50/67 , Y02T50/676 , Y10T428/12611 , Y10T428/12736 , Y10T428/12944 , Y10T428/24273 , Y10T428/24289 , Y10T428/24298 , Y10T428/24306
摘要： An actively cooled TBC bond coat wherein active convection cooling is provided through micro channels inside or adjacent to a bond coat layer applied to a substrate. The micro channels communicate directly with at least one cooling fluid supply contained within a turbine engine component, thereby providing direct and efficient cooling for the bond coat layer. Because the substrate is covered with an actively cooled bond coat layer, it will reduce the cooling requirement for the substrate, thus allowing the engine to run at higher operating temperature without the need for additional cooling air, achieving a better engine performance. In one form, the component includes a substrate having at least one substrate channel with a first and second end. At least one micro channel is in fluid communication with a plenum which in turn is in fluid communication with at least one substrate channel through an exit orifice in the substrate channel which is at a first end of the substrate channel. A second end of the substrate channel is in communication with a cooling fluid supply, for example, cooling circuits contained within the turbine engine component. The micro channel is located between the substrate surface and the outer gas flow path surface of the component.
摘要翻译：主动冷却的TBC粘合涂层，其中主动对流冷却通过施加到基底的粘合涂层的内部或附近的微通道提供。微通道与包含在涡轮发动机部件内的至少一个冷却流体供应件直接连通，从而为粘合涂层提供直接和有效的冷却。由于衬底被主动冷却的粘合涂层覆盖，所以它将降低衬底的冷却要求，从而允许发动机在更高的工作温度下运行，而不需要额外的冷却空气，实现更好的发动机性能。在一种形式中，组件包括具有至少一个具有第一和第二端的衬底通道的衬底。至少一个微通道与气室流体连通，该气室又与通过基板通道中位于衬底通道的第一端处的出口孔与至少一个衬底通道流体连通。衬底通道的第二端与冷却流体供应源连通，例如包含在涡轮发动机部件内的冷却回路。微通道位于基板表面和部件的外部气体流动通道表面之间。

7. 发明授权

US06528118B2 Process for creating structured porosity in thermal barrier coating 失效
公开(公告)号：US06528118B2
公开(公告)日：2003-03-04
申请号：US09777930
申请日：2001-02-06
申请人： Ching-Pang Lee , Ramgopal Darolia , Robert Edward Schafrik
发明人： Ching-Pang Lee , Ramgopal Darolia , Robert Edward Schafrik
IPC分类号： C23C1404
CPC分类号： C23C14/042 , C23C4/01 , C23C28/00 , Y02T50/67
摘要： A process for creating microgrooves within or adjacent to a TBC layer applied to a gas turbine engine component such as a blade or vane. The process includes the steps of applying a bond coat to the surface of the substrate. A wire mesh is placed a predetermined distance above the bond coat surface. With the wire mesh in position, about 0.002 inches of an inner TBC is applied over the bond coat. The wire in the wire mesh causes a shadow effect as the TBC is applied, so that there are variations in the thickness of the applied TBC, forming micro channels. The wire mesh is removed and an additional outer TBC layer is applied over the inner TBC layer, and the variations in thickness are bridged by the continued deposition of the columnar TBC over the inner TBC layer, forming the microgrooves.

8. 发明授权

US06499949B2 Turbine airfoil trailing edge with micro cooling channels 失效
标题翻译：涡轮机翼后缘带有微冷却通道
公开(公告)号：US06499949B2
公开(公告)日：2002-12-31
申请号：US09818385
申请日：2001-03-27
申请人： Robert Edward Schafrik , Ramgopal Darolia , Ching-Pang Lee
发明人： Robert Edward Schafrik , Ramgopal Darolia , Ching-Pang Lee
IPC分类号： F01D518
CPC分类号： F01D5/187 , F05D2230/14 , Y02T50/67 , Y02T50/673 , Y02T50/676 , Y10T29/4932
摘要： The present invention provides active convection cooling through micro channels within or adjacent to a bond coat layer applied to the trailing edge of a turbine engine high pressure airfoil. When placed adjacent to or within a porous TBC, the micro channels additionally provide transpiration cooling through the porous TBC. The micro channels communicate directly with at least one cooling circuit contained within the airfoil from which they receive cooling air, thereby providing direct and efficient cooling for the bond coat layer. Because the substrate includes an actively cooled flow path surface region that can reduce the cooling requirement for the substrate, the engine can run at a higher firing temperature without the need for additional cooling air, achieving a better, more efficient engine performance. In one embodiment, a metallic bond coat is added to an airfoil with pressure side bleed film cooling slots. The bond coat is grooved such that the grooves are structured, with at least one structured micro groove communicating with at least one cooling fluid supply contained within the airfoil. A TBC layer is applied, using a shadowing technique, over the structured grooves, resulting in the formation of hollow micro channels for the transport of the cooling fluid. In different embodiments, the location of the structured grooves, hence, the resulting micro channels are placed within the airfoil substrate at the substrate/bond coat interface or within the TBC layer.
摘要翻译：本发明提供了在施加到涡轮发动机高压翼型件的后缘上的粘合涂层之内或附近的微通道的主动对流冷却。当放置在多孔TBC附近或内部时，微通道另外通过多孔TBC提供蒸腾冷却。微通道与包含在翼片内的至少一个冷却回路直接连通，从而从它们接收冷却空气，从而为粘结涂层提供直接和有效的冷却。因为基板包括可以降低基板的冷却要求的主动冷却的流动路径表面区域，所以发动机可以在更高的点火温度下运行，而不需要额外的冷却空气，实现更好，更有效的发动机性能。在一个实施例中，将金属粘合涂层加入到具有压力侧泄放膜冷却槽的翼型件上。接合涂层是开槽的，使得凹槽被构造成具有与包含在翼型内的至少一个冷却流体供应源连通的至少一个结构化微槽。使用阴影技术在结构化凹槽上施加TBC层，导致形成用于输送冷却流体的中空微通道。在不同的实施例中，结构化凹槽的位置，因此，所得到的微通道在衬底/粘结涂层界面处或在TBC层内被放置在翼型衬底内。

9. 发明授权

US06461107B1 Turbine blade tip having thermal barrier coating-formed micro cooling channels 失效
标题翻译：涡轮叶片尖端具有形成热障涂层的微冷却通道
公开(公告)号：US06461107B1
公开(公告)日：2002-10-08
申请号：US09818312
申请日：2001-03-27
申请人： Ching-Pang Lee , Ramgopal Darolia , Robert Edward Schafrik
发明人： Ching-Pang Lee , Ramgopal Darolia , Robert Edward Schafrik
IPC分类号： F01D518
CPC分类号： F01D5/288 , F01D5/183 , F01D5/187 , F01D5/20 , Y02T50/67 , Y02T50/673 , Y02T50/676 , Y10T29/49318
摘要： The present invention provides for cooling the squealer tip region of a high pressure turbine blade used in a gas turbine engine comprising coating the squealer tip with a metallic bond coat. Micro grooves oriented in the radial direction are fabricated into the airfoil on the interior surface of the squealer tip above and substantially perpendicular to the tip cap. A micro groove oriented in the axial direction is fabricated along the joint corner between the squealer tip side wall and the, tip cap to connect and act as a plenum with all of the micro grooves oriented in the radial direction. Tip cap cooling holes are drilled through the tip cap and connected to the micro groove that ultimately forms a plenum. TBC ceramic is then deposited on both blade external surfaces and the tip cavity, forming micro channels from micro grooves as a result of self shadowing. In this manner, cooling fluid passes from a cooling fluid source through the tip cap holes and into the plenum created by the micro channel, subsequently passing into the micro channels that are oriented in the radial direction. Cooling fluid is thereby directed through the micro channels to cool the squealer, exiting in the vicinity of the tip. Since the TBC is porous, some of the cooling fluid will also flow through the TBC to provide transpiration cooling. The present invention further comprises both the cooled blade and squealer tip region formed by the foregoing methods and the blade and squealer tip with the micro channels for cooling the squealer tip.
摘要翻译：本发明提供了用于冷却在燃气涡轮发动机中使用的高压涡轮机叶片的鸣响器尖端区域，其包括用金属粘结涂层涂覆尖叫尖端。在径向方向上定向的微槽在尖端顶部的内表面上被制造成位于尖端顶部的内表面上且基本上垂直于顶盖。沿着轴向定向的微槽沿着尖叫尖端侧壁和尖端盖之间的接合角制造，以连接并充当具有沿径向方向定向的所有微槽的集气室。顶盖冷却孔穿过尖端盖并连接到最终形成增压室的微槽。然后将TBC陶瓷沉积在两个叶片外表面和尖端腔上，由于自身阴影而从微槽形成微通道。以这种方式，冷却流体从冷却流体源通过顶盖孔并进入由微通道产生的增压室，随后进入沿径向定向的微通道。因此，冷却流体被引导通过微通道，以冷却在尖端附近离开的鸣叫器。由于TBC是多孔的，一些冷却流体也将流经TBC以提供蒸发冷却。本发明还包括通过上述方法形成的冷却刀片和鸣响器尖端区域以及具有用于冷却鸣叫器尖端的微通道的刀片和尖叫尖端。

10. 发明申请

US20060099080A1 Methods and apparatus for cooling gas turbine engine components 失效
标题翻译：用于冷却燃气涡轮发动机部件的方法和装置
公开(公告)号：US20060099080A1
公开(公告)日：2006-05-11
申请号：US10984292
申请日：2004-11-09
申请人： Ching-Pang Lee , Ronald Bunker , Harvey Maclin , Ramgopal Darolia
发明人： Ching-Pang Lee , Ronald Bunker , Harvey Maclin , Ramgopal Darolia
IPC分类号： F03B3/12
CPC分类号： F23R3/002 , F01D5/183 , F01D5/186 , F01D5/288 , F05D2230/90 , F05D2300/611 , F23R2900/03041
摘要： A method of cooling a gas turbine engine component having a perforate metal wall includes providing a plurality of pores in the wall, wherein the pores extend substantially perpendicularly through the wall, and wherein the pores are covered and sealed closed at first ends thereof by a thermal barrier coating disposed over a first surface of the wall, and providing a plurality of film cooling holes in the wall, wherein the holes extend substantially perpendicularly through the wall and the thermal barrier coating. The method also includes providing cooling fluid to the plurality of pores and the plurality of film cooling holes along a second surface of the wall, channeling the cooling fluid through the pores for back side cooling an inner surface of the thermal barrier coating, and channeling the cooling fluid through the holes for film cooling an outer surface of the thermal barrier coating.
摘要翻译：一种冷却具有穿孔金属壁的燃气涡轮发动机部件的方法包括在壁中提供多个孔，其中孔基本上垂直延伸穿过壁，并且其中孔在其第一端被热封屏障涂层，其设置在所述壁的第一表面上，并且在所述壁中提供多个膜冷却孔，其中所述孔基本上垂直延伸穿过所述壁和所述热障涂层。该方法还包括沿着壁的第二表面向多个孔和多个膜冷却孔提供冷却流体，将冷却流体引导通过孔，用于背面冷却热障涂层的内表面，并且引导冷却流体通过孔，用于薄膜冷却隔热涂层的外表面。

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