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    Turbine blade tip with vortex generators
    2.
    发明授权
    Turbine blade tip with vortex generators 失效
    涡轮叶片尖端与涡流发生器

    公开(公告)号：US08690536B2

    公开(公告)日：2014-04-08

    申请号：US12892515

    申请日：2010-09-28

    申请人： Alexander R. Beeck ,  Glenn E. Brown

    发明人： Alexander R. Beeck ,  Glenn E. Brown

    IPC分类号： F01D5/20 ,  F01D5/08

    CPC分类号： F01D5/20 ,  F01D5/187 ,  F05D2210/33 ,  F05D2240/127 ,  F05D2250/21 ,  F05D2260/2212

    摘要： A turbine blade for a turbine engine having a tip with one or more vortex generators for reducing tip leakage during operation of the turbine engine. The vortex generators may extend radially outward from the radially outer surface of the tip wall. The vortex generator may be positioned between a rib extending radially outward from the radially outer surface of the tip wall and an intersection between the outer surface of the tip wall and an outer surface on the pressure side. The vortex generators may include a base and three sides forming a triangular point with a first side having a larger surface are than second and third sides. One or more film cooling holes may be formed in the tip wall to provide cooling air to the tip. In one embodiment, film cooling holes may be positioned in one or more vortex generators.

    摘要翻译： 一种用于涡轮发动机的涡轮叶片，其具有具有一个或多个涡流发生器的尖端，用于在涡轮发动机的运行期间减少尖端泄漏。 涡流发生器可以从尖端壁的径向外表面径向向外延伸。 涡流发生器可以位于从尖端壁的径向外表面径向向外延伸的肋与尖端壁的外表面与压力侧的外表面之间的交叉点之间。 涡流发生器可以包括基部，并且形成具有比第二和第三侧更大表面的第一侧的三角形点的三个边。 可以在尖端壁中形成一个或多个薄膜冷却孔，以向尖端提供冷却空气。 在一个实施例中，膜冷却孔可以定位在一个或多个涡流发生器中。

    SERPENTINE COOLING CIRCUIT WITH T-SHAPED PARTITIONS IN A TURBINE AIRFOIL
    3.
    发明申请
    SERPENTINE COOLING CIRCUIT WITH T-SHAPED PARTITIONS IN A TURBINE AIRFOIL 有权
    在涡轮机空气中的T形成型的SERPENTINE冷却回路

    公开(公告)号：US20120269648A1

    公开(公告)日：2012-10-25

    申请号：US13092303

    申请日：2011-04-22

    申请人： Ching-Pang Lee

    发明人： Ching-Pang Lee

    IPC分类号： F01D5/18

    CPC分类号： F01D5/187 ,  F05D2210/33 ,  F05D2250/185 ,  F05D2260/2212

    摘要： A serpentine cooling circuit (AFT) in a turbine airfoil (34A) starting from a radial feed channel (C1), and progressing axially (65) in alternating tangential directions through interconnected channels (C1, C2, C3) formed between partitions (T1, T2, J1). At least one of the partitions (T1, T2) has a T-shaped transverse section, with a base portion (67) extending from a suction or pressure side wall (64, 62) of the airfoil, and a crossing portion (68, 69) parallel to, and not directly attached to, the opposite pressure or suction side wall (62, 64). Each crossing portion bounds a near-wall passage (N1, N2) adjacent to the opposite pressure or suction side wall (62, 64). Each near-wall passage may have a smaller flow aperture area than one, or each, of two adjacent connected channels (C1, C2, C3). The serpentine circuit (AFT) may follow a forward cooling circuit (FWD) in the airfoil (34A).

    摘要翻译： 涡轮机翼型（34A）中的蛇形冷却回路（AFT），其从径向进料通道（C1）开始，并且沿交替的切线方向通过形成在分隔件（T1，C2）之间的互连通道（C1，C2，C3） T2，J1）。 至少一个隔板（T1，T2）具有T形横截面，其中从翼型件的吸力或压力侧壁（64,62）延伸的基部（67）和交叉部分 69）平行于并且不直接附接到相对的压力或吸力侧壁（62,64）。 每个交叉部分限定与相对的压力或吸力侧壁（62,64）相邻的近壁通道（N1，N2）。 每个近壁通道可以具有比两个相邻连接的通道（C1，C2，C3）中的一个或每个通道更小的流通孔面积。 蛇形回路（AFT）可以跟随机翼（34A）中的向前冷却回路（FWD）。

    Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles
    6.
    发明授权
    Turbine airfoil vane with an impingement insert having a plurality of impingement nozzles 有权
    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。

    Dual-Vortical-Flow Hybrid Rocket Engine
    8.
    发明申请
    Dual-Vortical-Flow Hybrid Rocket Engine 有权
    双涡流混合火箭发动机

    公开(公告)号：US20140352276A1

    公开(公告)日：2014-12-04

    申请号：US14070881

    申请日：2013-11-04

    申请人： National Chiao Tung University ,  National Applied Research Laboratories

    发明人： Yen-Sen Chen ,  Men-Zen Wu ,  Jong-Shinn Wu ,  Alfred Lai ,  Jhe-Wei Lin ,  Tzu-Hao Chou

    IPC分类号： F02K9/00

    CPC分类号： F02K9/72 ,  F05D2210/33 ,  F05D2240/35

    摘要： The present invention discloses a dual-vortical-flow hybrid rocket engine, including a main body and a nozzle communicating with an end of the main body. The main body includes a plurality of disk-like combustion chambers arranged longitudinally, and a central combustion chamber formed along the axial portion and communicating the disk-like combustion chambers. Each of the disk-like combustion chambers is provided with a plurality of oxidizer injection nozzles at its inner circumference surface. Inside the disk-like combustion chambers, the oxidizer is injected in nearly the tangent directions of the circumference, and the injection directions are opposite for the neighboring disk-like combustion chambers, which creates vortical flows with opposite rotating directions so as to increase the total residence time of the combustion reactions of the oxidizer and the solid-state fuel in the disk-like combustion chambers of the present invention.

    摘要翻译： 本发明公开了一种双涡旋混合火箭发动机，其包括主体和与主体的端部连通的喷嘴。 主体包括纵向布置的多个盘形燃烧室和沿轴向部分形成并连通圆盘状燃烧室的中心燃烧室。 每个盘状燃烧室在其内圆周表面上设置有多个氧化剂注入喷嘴。 在盘状燃烧室内部，氧化剂沿着圆周的切线方向喷射，并且相邻的盘状燃烧室的喷射方向相反，这形成相反旋转方向的涡旋流，从而增加总量 本发明的盘状燃烧室中的氧化剂和固体燃料的燃烧反应的停留时间。

    TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY OF IMPINGEMENT NOZZLES
    9.
    发明申请
    TURBINE AIRFOIL VANE WITH AN IMPINGEMENT INSERT HAVING A PLURALITY OF IMPINGEMENT NOZZLES 有权
    具有多个冲击喷嘴的涡旋空气涡街风扇

    公开(公告)号：US20140105726A1

    公开(公告)日：2014-04-17

    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

    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。

    EXHAUST SYSTEM FOR GAS TURBINES
    10.
    发明申请
    EXHAUST SYSTEM FOR GAS TURBINES 审中-公开
    排气系统用于气体涡轮机

    公开(公告)号：US20130327052A1

    公开(公告)日：2013-12-12

    申请号：US13916396

    申请日：2013-06-12

    申请人： GENALTA POWER INC.

    发明人： Chris O'NEILL ,  Dobromir FILIP ,  Joseph KISS ,  Graham ILLINGWORTH

    IPC分类号： F02C6/18 ,  F02C7/141

    CPC分类号： F02C6/18 ,  F01D25/30 ,  F02C6/00 ,  F02C7/141 ,  F05D2210/33

    摘要： An exhaust system is provided for mitigating condensate formation in a common exhaust stack and for effecting improved heat transfer. Reduced condensate formation and improved heat transfer is achieved by inducing non-laminar flow through the common exhaust stack and a heat exchanger operatively coupled to the common exhaust stack. Heat transfer is further improved by dew point control. Non-laminar flow is induced by connecting more than one gas turbine to the common exhaust stack through non-laminar flow inducing arrangements. The various coupling arrangements also add structural rigidity to the common exhaust stack for increased stack height and improved plume dispersion.

    摘要翻译： 提供了一种排气系统，用于减轻公共排气叠层中的冷凝物形成并实现改进的热传递。 通过引导非层流流经公共排气叠层和可操作地连接到公共排气叠层的热交换器来实现冷凝物形成减少和热传递的改善。 通过露点控制进一步提高了热传递。 通过非层流引导装置将多于一个燃气轮机连接到公共排气叠层引起非层流。 各种联接布置还增加了公共排气叠层的结构刚度，以增加堆叠高度和改进的羽流分散。