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1. 发明授权

US08668453B2 Cooling system having reduced mass pin fins for components in a gas turbine engine 有权
标题翻译：具有减少用于燃气涡轮发动机中的部件的散热片的冷却系统
公开(公告)号：US08668453B2
公开(公告)日：2014-03-11
申请号：US13027323
申请日：2011-02-15
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra
IPC分类号： F01D5/08
CPC分类号： F01D5/187 , F05D2210/33 , F05D2240/127 , F05D2250/32 , F05D2250/70 , F05D2260/2212
摘要： A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil.
摘要翻译：公开了一种具有用于燃气涡轮发动机的具有减小质量的一个或多个销翅片的冷却系统。冷却系统可以包括限定冷却系统的至少一部分的一个或多个第一表面。销翅片可以从限定冷却系统的表面延伸并且可以具有大致平行于表面的非圆形截面，并且横截面的外表面的至少一部分形成至少四分之一圆。销翅片的下游侧可以具有空腔以减小质量，从而形成更有效的涡轮机翼型件。

2. 发明申请

US20120207591A1 COOLING SYSTEM HAVING REDUCED MASS PIN FINS FOR COMPONENTS IN A GAS TURBINE ENGINE 有权
标题翻译：具有减少燃油涡轮发动机组件的减少质量引脚的冷却系统
公开(公告)号：US20120207591A1
公开(公告)日：2012-08-16
申请号：US13027323
申请日：2011-02-15
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra
IPC分类号： F01D5/14
CPC分类号： F01D5/187 , F05D2210/33 , F05D2240/127 , F05D2250/32 , F05D2250/70 , F05D2260/2212
摘要： A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil.
摘要翻译：公开了一种具有用于燃气涡轮发动机的具有减小质量的一个或多个销翅片的冷却系统。冷却系统可以包括限定冷却系统的至少一部分的一个或多个第一表面。销翅片可以从限定冷却系统的表面延伸并且可以具有大致平行于表面的非圆形截面，并且横截面的外表面的至少一部分形成至少四分之一圆。销翅片的下游侧可以具有空腔以减小质量，从而形成更有效的涡轮机翼型件。

3. 发明申请

US20110033312A1 COMPOUND COOLING FLOW TURBULATOR FOR TURBINE COMPONENT 有权
标题翻译：涡轮组件的复合冷却流量涡轮机
公开(公告)号：US20110033312A1
公开(公告)日：2011-02-10
申请号：US12884464
申请日：2010-09-17
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph
IPC分类号： F01D5/18
CPC分类号： F01D5/18 , F01D5/187 , F05D2250/181 , F05D2250/60 , F05D2250/611 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2260/2212 , F05D2260/22141
摘要： Multi-scale turbulation features, including first turbulators (46, 48) on a cooling surface (44), and smaller turbulators (52, 54, 58, 62) on the first turbulators. The first turbulators may be formed between larger turbulators (50). The first turbulators may be alternating ridges (46) and valleys (48). The smaller turbulators may be concave surface features such as dimples (62) and grooves (54), and/or convex surface features such as bumps (58) and smaller ridges (52). An embodiment with convex turbulators (52, 58) in the valleys (48) and concave turbulators (54, 62) on the ridges (46) increases the cooling surface area, reduces boundary layer separation, avoids coolant shadowing and stagnation, and reduces component mass.
摘要翻译：多尺度紊流特征，包括在冷却表面（44）上的第一湍流器（46,48）和在第一湍流器上的较小湍流器（52,54,58,62）。可以在较大的湍流器（50）之间形成第一湍流器。第一个紊流器可以是交替的脊（46）和谷（48）。较小的湍流器可以是诸如凹坑（62）和凹槽（54）的凹表面特征，和/或诸如凸块（58）和较小脊（52）的凸表面特征。在脊（46）上的谷（48）和凹形湍流器（54,62）中具有凸起的湍流器（52,58）的实施例增加了冷却表面面积，减少了边界层分离，避免了冷却剂遮蔽和停滞，并且减少了部件质量

4. 发明申请

US20180318919A1 TURBINE AIRFOIL COOLING SYSTEM WITH LEADING EDGE IMPINGEMENT COOLING SYSTEM TURBINE BLADE INVESTMENT CASTING USING FILM HOLE PROTRUSIONS FOR INTEGRAL WALL THICKNESS CONTROL 审中-公开
公开(公告)号：US20180318919A1
公开(公告)日：2018-11-08
申请号：US16036611
申请日：2018-07-16
申请人： Ching-Pang Lee , Nan Jiang
发明人： Ching-Pang Lee , Nan Jiang
IPC分类号： B22D25/02 , F01D5/18 , B22C9/04 , B22C9/10 , B22C7/02 , B22D29/00 , B22C21/14 , B22C9/24
CPC分类号： B22D25/02 , B22C7/02 , B22C9/04 , B22C9/043 , B22C9/10 , B22C9/108 , B22C9/24 , B22C21/14 , B22D29/001 , F01D5/186 , F05D2220/32 , F05D2230/211
摘要： 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.

5. 发明授权

US08894367B2 Compound cooling flow turbulator for turbine component 有权
标题翻译：用于涡轮机部件的复合式冷却流动紊流器
公开(公告)号：US08894367B2
公开(公告)日：2014-11-25
申请号：US12884464
申请日：2010-09-17
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph
IPC分类号： F01D5/18
CPC分类号： F01D5/18 , F01D5/187 , F05D2250/181 , F05D2250/60 , F05D2250/611 , F05D2250/70 , F05D2250/711 , F05D2250/712 , F05D2260/2212 , F05D2260/22141
摘要： Multi-scale turbulation features, including first turbulators (46, 48) on a cooling surface (44), and smaller turbulators (52, 54, 58, 62) on the first turbulators. The first turbulators may be formed between larger turbulators (50). The first turbulators may be alternating ridges (46) and valleys (48). The smaller turbulators may be concave surface features such as dimples (62) and grooves (54), and/or convex surface features such as bumps (58) and smaller ridges (52). An embodiment with convex turbulators (52, 58) in the valleys (48) and concave turbulators (54, 62) on the ridges (46) increases the cooling surface area, reduces boundary layer separation, avoids coolant shadowing and stagnation, and reduces component mass.
摘要翻译：多尺度紊流特征，包括在冷却表面（44）上的第一湍流器（46,48）和在第一湍流器上的较小湍流器（52,54,58,62）。可以在较大的湍流器（50）之间形成第一湍流器。第一个紊流器可以是交替的脊（46）和谷（48）。较小的湍流器可以是诸如凹坑（62）和凹槽（54）的凹表面特征，和/或诸如凸块（58）和较小脊（52）的凸表面特征。在脊（46）上的谷（48）和凹形湍流器（54,62）中具有凸起的湍流器（52,58）的实施例增加了冷却表面面积，减少了边界层分离，避免了冷却剂遮蔽和停滞，并且减少了部件质量

6. 发明授权

US09022736B2 Integrated axial and tangential serpentine cooling circuit in a turbine airfoil 有权
标题翻译：涡轮机翼中集成的轴向和切向蛇形冷却回路
公开(公告)号：US09022736B2
公开(公告)日：2015-05-05
申请号：US13027333
申请日：2011-02-15
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph , John P. Dalton
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph , John P. Dalton
IPC分类号： F01D5/18
CPC分类号： F01D5/188 , F01D5/186 , F01D5/187 , F05D2250/185 , F05D2260/2212
摘要： A continuous serpentine cooling circuit forming a progression of radial passages (44, 45, 46, 47A, 48A) between pressure and suction side walls (52, 54) in a MID region of a turbine airfoil (24). The circuit progresses first axially, then tangentially, ending in a last radial passage (48A) adjacent to the suction side (54) and not adjacent to the pressure side (52). The passages of the axial progression (44, 45, 46) may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (47A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (47A, 48A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.
摘要翻译：连续的蛇形冷却回路，形成在涡轮机翼型件（24）的MID区域中的压力和吸力侧壁（52,54）之间的径向通道（44,45,46,47A，48A）的进展。电路首先轴向前进，然后切向地结束于邻近吸力侧（54）并且不与压力侧（52）相邻的最后一个径向通道（48A）。轴向行进（44,45,46）的通道可以与翼型件的压力和吸力侧壁相邻。最后一个径向通道（47A）的下一个可以与压力侧壁相邻并且不与吸力侧壁相邻。最后两个径向通道（47A，48A）可以沿着压力和吸力侧壁分别比在宽度方向上更长，从而在这些热壁的内部提供增加的直接冷却表面积。

7. 发明申请

US20120207614A1 INTEGRATED AXIAL AND TANGENTIAL SERPENTINE COOLING CIRCUIT IN A TURBINE AIRFOIL 有权
标题翻译：涡轮机空气中的集成式轴向和螺旋式冷却回路
公开(公告)号：US20120207614A1
公开(公告)日：2012-08-16
申请号：US13027333
申请日：2011-02-15
申请人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph , John P. Dalton
发明人： Ching-Pang Lee , Nan Jiang , John J. Marra , Ronald J. Rudolph , John P. Dalton
IPC分类号： F01D5/18
CPC分类号： F01D5/188 , F01D5/186 , F01D5/187 , F05D2250/185 , F05D2260/2212
摘要： A continuous serpentine cooling circuit forming a progression of radial passages (44, 45, 46, 47A, 48A) between pressure and suction side walls (52, 54) in a MID region of a turbine airfoil (24). The circuit progresses first axially, then tangentially, ending in a last radial passage (48A) adjacent to the suction side (54) and not adjacent to the pressure side (52). The passages of the axial progression (44, 45, 46) may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (47A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (47A, 48A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.
摘要翻译：连续的蛇形冷却回路，形成在涡轮机翼型件（24）的MID区域中的压力和吸力侧壁（52,54）之间的径向通道（44,45,46,47A，48A）的进展。电路首先轴向前进，然后切向地结束于邻近吸力侧（54）并且不与压力侧（52）相邻的最后一个径向通道（48A）。轴向行进（44,45,46）的通道可以与翼型件的压力和吸力侧壁相邻。最后一个径向通道（47A）的下一个可以与压力侧壁相邻并且不与吸力侧壁相邻。最后两个径向通道（47A，48A）可以沿着压力和吸力侧壁分别比在宽度方向上更长，从而在这些热壁的内部提供增加的直接冷却表面积。

8. 发明授权

US09630277B2 Airfoil having built-up surface with embedded cooling passage 有权
公开(公告)号：US09630277B2
公开(公告)日：2017-04-25
申请号：US12723880
申请日：2010-03-15
申请人： Ching-Pang Lee , Mrinal Munshi , Gm S. Azad , Jae Y. Um
发明人： Ching-Pang Lee , Mrinal Munshi , Gm S. Azad , Jae Y. Um
IPC分类号： F02C7/12 , B23P6/00 , F01D5/00 , F01D5/14 , F01D5/18
CPC分类号： B23P6/007 , F01D5/005 , F01D5/145 , F01D5/186 , F05D2230/30 , F05D2230/80 , F05D2240/81 , Y10T29/49318
摘要： A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform.

9. 发明授权

US09347324B2 Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles 有权
标题翻译：具有冲击插入件的涡轮机翼型叶片具有多个冲击喷嘴
公开(公告)号：US09347324B2
公开(公告)日：2016-05-24
申请号：US14133773
申请日：2013-12-19
申请人： Ching-Pang Lee
发明人： Ching-Pang Lee
IPC分类号： F01D5/18 , F01D9/04
CPC分类号： F01D5/188 , F01D9/04 , F05D2210/33 , F05D2240/12 , F05D2240/127 , F05D2260/2212
摘要： A turbine vane includes a generally elongated hollow airfoil and a cooling system. The cooling system is positioned within the airfoil and includes a cooling chamber and an impingement insert positioned in the cooling chamber. The impingement insert and an inner surface of an outer wall of the airfoil define a cooling channel therebetween. The impingement insert includes a plurality of impingement nozzles extending toward the inner surface of the outer wall and a plurality of impingement orifices. At least one of the impingement orifices is arranged in a non-aligned pattern with respect to at least one adjacent impingement orifice such that cooling fluid passing out of the at least one impingement orifice does not directly flow into a centerline of a cooling fluid flowpath of cooling fluid passing out of the at least one adjacent impingement orifice.
摘要翻译：涡轮叶片包括大致细长的中空翼型件和冷却系统。冷却系统定位在翼型内，并且包括冷却室和位于冷却室中的冲击插入件。冲击插入件和翼型件的外壁的内表面在它们之间限定冷却通道。冲击插入件包括朝向外壁的内表面延伸的多个冲击喷嘴和多个冲击孔。冲击孔中的至少一个相对于至少一个相邻的冲击孔布置成非对准图案，使得流出至少一个冲击孔的冷却流体不直接流入冷却流体流动路径的中心线从至少一个相邻的冲击孔流出的冷却流体。

10. 发明授权

US09206904B2 Seal including flexible seal strips 有权
标题翻译：密封包括柔性密封条
公开(公告)号：US09206904B2
公开(公告)日：2015-12-08
申请号：US12832141
申请日：2010-07-08
申请人： Ching-Pang Lee , Vincent P. Laurello , Chander Prakash , Kok-Mun Tham
发明人： Ching-Pang Lee , Vincent P. Laurello , Chander Prakash , Kok-Mun Tham
IPC分类号： F16J15/32 , F01D11/02
CPC分类号： F16J15/3292 , F01D11/02 , F05D2240/57
摘要： A seal member for effecting a seal preventing fluid flow in an axial direction through an annular space formed between two relatively moving components including a rotatable shaft and a stator structure. The seal member includes a plurality of flexible seal strips. Each seal strip includes a planar plate extending radially through the annular space and having a radially outer end supported to the stator structure and a radially inner end defining a tip portion extending widthwise in the axial direction engaged in sliding contact with a peripheral surface of the rotatable shaft. At least one of the seal strips includes a plurality of perforations extending through the seal strip and located between a leading edge and a trailing edge of the seal strip for effecting an increased flexibility of the seal strip adjacent to the tip portion.
摘要翻译：一种用于实现密封的密封件，其防止流体沿轴向流动通过形成在包括可旋转轴和定子结构的两个相对移动部件之间的环形空间。密封构件包括多个柔性密封条。每个密封条包括平面板，该平面板径向地延伸穿过环形空间并且具有支撑到定子结构的径向外端，以及限定沿轴向方向沿宽度方向延伸的尖端部分的径向内端，该尖端部分与可旋转的轴。密封条中的至少一个包括延伸穿过密封条并且位于密封条的前缘和后缘之间的多个穿孔，用于实现与尖端部分相邻的密封条的增加的柔性。

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