公开(公告)号：US09091180B2

公开(公告)日：2015-07-28

申请号：US13552708

申请日：2012-07-19

申请人： David J. Wiebe ,  Bruce L. Smith ,  Matthew D. Montgomery

发明人： David J. Wiebe ,  Bruce L. Smith ,  Matthew D. Montgomery

IPC分类号： F01D25/00 ,  F01D9/06 ,  F01D5/18 ,  F01D5/14

CPC分类号： F01D25/00 ,  F01D5/145 ,  F01D5/18 ,  F01D9/041 ,  F01D9/06 ,  F05D2240/80 ,  Y10T29/49234

摘要： An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge.

摘要翻译： 一种翼型组件，其包括端壁和从所述气流通道延伸的翼型件。 端壁包括上游和下游边缘，并且限定在具有在平台结构的厚度方向上径向延伸的前表面的平台结构上。 至少一个流体喷射通道在从端壁的上游边缘向下游边缘的方向上延伸穿过平台结构。 流体注入通道具有限定在端壁处的出口开口和与加压流体源流体连通的入口开口。 流体喷射通道相对于端壁的平面以较小的角度延伸，其中流体喷射通道限定通过前表面和端壁的通道轴线，以实现出口和翼型前缘之间的边界层的通电。

公开(公告)号：US20140023483A1

公开(公告)日：2014-01-23

申请号：US13552708

申请日：2012-07-19

申请人： David J. Wiebe ,  Bruce L. Smith ,  Matthew D. Montgomery

发明人： David J. Wiebe ,  Bruce L. Smith ,  Matthew D. Montgomery

IPC分类号： F01D25/00 ,  F01D5/18 ,  F01D9/06

CPC分类号： F01D25/00 ,  F01D5/145 ,  F01D5/18 ,  F01D9/041 ,  F01D9/06 ,  F05D2240/80 ,  Y10T29/49234

摘要： An airfoil assembly including an endwall and an airfoil extending from the into a gas flow path. The endwall includes upstream and downstream edges, and is defined on a platform structure having a front surface extending radially in a direction of a thickness of the platform structure. At least one fluid injection passage extends through the platform structure in a direction from the upstream edge toward the downstream edge of the endwall. The fluid injection passage has an outlet opening defined at the endwall and an inlet opening in fluid communication with a pressurized fluid source. The fluid injection passage extends at a shallow angle relative to a plane of the endwall wherein the fluid injection passage defines a passage axis passing through the front surface and the endwall for effecting energization of a boundary layer between the outlet opening and the airfoil leading edge.

摘要翻译： 一种翼型组件，其包括端壁和从所述气流通道延伸的翼型件。 端壁包括上游和下游边缘，并且限定在具有在平台结构的厚度方向上径向延伸的前表面的平台结构上。 至少一个流体喷射通道在从端壁的上游边缘向下游边缘的方向上延伸穿过平台结构。 流体注入通道具有限定在端壁处的出口开口和与加压流体源流体连通的入口开口。 流体喷射通道相对于端壁的平面以较小的角度延伸，其中流体喷射通道限定通过前表面和端壁的通道轴线，以实现出口和翼型前缘之间的边界层的通电。

公开(公告)号：US08585356B2

公开(公告)日：2013-11-19

申请号：US12729380

申请日：2010-03-23

申请人： David J. Wiebe ,  Matthew D. Montgomery

发明人： David J. Wiebe ,  Matthew D. Montgomery

IPC分类号： F01D11/20

CPC分类号： F01D11/10 ,  F01D11/20 ,  F05D2270/172

摘要： An electrode (54) in the tip (31) of a turbine or compressor blade (30), and a series of electrodes (68) in a shroud (36, 64) that surrounds a rotation path (33) of the blade tip. As the blade tip reaches each shroud electrode, a controller (74) activates an electrical potential between them that generates a plasma-induced gas flow (76) directed toward the pressure side (PS) of the airfoil. The plasma creates a seal between the blade tip and the shroud, and induces a gas flow that opposes a leakage gas flow (52) from the pressure side to the suction side (SS) of the blade over the blade tip (31).

摘要翻译： 涡轮或压缩机叶片（30）的尖端（31）中的电极（54）和围绕叶片尖端的旋转路径（33）的护罩（36,64）中的一系列电极（68）。 当叶片尖端到达每个护罩电极时，控制器（74）激活它们之间的电位，其产生指向翼型的压力侧（PS）的等离子体引起的气流（76）。 等离子体在叶片尖端和护罩之间产生密封，并且引起气体流动，该气流与来自叶片尖端（31）的叶片的压力侧到吸力侧（SS）的泄漏气体流（52）相对。

公开(公告)号：US20110236182A1

公开(公告)日：2011-09-29

申请号：US12729380

申请日：2010-03-23

申请人： David J. Wiebe ,  Matthew D. Montgomery

发明人： David J. Wiebe ,  Matthew D. Montgomery

IPC分类号： F01D15/00 ,  F01D11/10

CPC分类号： F01D11/10 ,  F01D11/20 ,  F05D2270/172

摘要： An electrode (54) in the tip (31) of a turbine or compressor blade (30), and a series of electrodes (68) in a shroud (36, 64) that surrounds a rotation path (33) of the blade tip. As the blade tip reaches each shroud electrode, a controller (74) activates an electrical potential between them that generates a plasma-induced gas flow (76) directed toward the pressure side (PS) of the airfoil. The plasma creates a seal between the blade tip and the shroud, and induces a gas flow that opposes a leakage gas flow (52) from the pressure side to the suction side (SS) of the blade over the blade tip (31).

摘要翻译： 涡轮或压缩机叶片（30）的尖端（31）中的电极（54）和围绕叶片尖端的旋转路径（33）的护罩（36,64）中的一系列电极（68）。 当叶片尖端到达每个护罩电极时，控制器（74）激活它们之间的电位，其产生指向翼型的压力侧（PS）的等离子体引起的气流（76）。 等离子体在叶片尖端和护罩之间产生密封，并且引起气体流动，该气流与来自叶片尖端（31）的叶片的压力侧到吸力侧（SS）的泄漏气体流（52）相对。

公开(公告)号：US20150000283A1

公开(公告)日：2015-01-01

申请号：US13928476

申请日：2013-06-27

申请人： Stephen A. Ramier ,  David J. Wiebe ,  Lawrence P. Budnick ,  Mark L. Adamson ,  Daniel W. Garan ,  Robert H. Bartley ,  Yadollah Naghian

发明人： Stephen A. Ramier ,  David J. Wiebe ,  Lawrence P. Budnick ,  Mark L. Adamson ,  Daniel W. Garan ,  Robert H. Bartley ,  Yadollah Naghian

IPC分类号： F23R3/28

CPC分类号： F23R3/283 ,  F16L19/061 ,  F23R3/286 ,  F23R3/34 ,  F23R3/346 ,  F23R3/54 ,  F23R3/60

摘要： A combustor apparatus defining a combustion zone where air and fuel are burned to create high temperature combustion products. The combustor apparatus comprises an outer wall including a fuel inlet opening for receiving a fuel feed pipe. A coupling assembly is engaged with the fuel feed pipe at the fuel inlet opening to attach the fuel feed pipe to the outer wall. A fuel injection system is located in the interior volume of the outer wall and comprises fuel supply structure including a fuel feed block having a fuel intake passage aligned with the outlet portion of the fuel feed pipe. A coupling fastener is engaged against an exterior outer face of the fuel feed block to create a sealed coupling for containing fuel passing from the fuel feed pipe into the fuel feed block, and to secure the fuel feed block relative to the coupling assembly.

摘要翻译： 一种限定燃烧区域的燃烧器装置，其中燃烧空气和燃料以产生高温燃烧产物。 燃烧器装置包括外壁，其包括用于接收燃料供给管的燃料入口。 联接组件在燃料入口处与燃料供给管接合以将燃料供给管附接到外壁。 燃料喷射系统位于外壁的内部容积中，并且包括燃料供给结构，其包括具有与燃料供给管的出口部分对准的燃料进入通道的燃料供给块。 联接紧固件抵靠燃料供给块的外部外表面接合以产生密封联接件，用于容纳从燃料供给管道进入燃料供给块的燃料，并且相对于联接组件固定燃料供给块。

公开(公告)号：US08668454B2

公开(公告)日：2014-03-11

申请号：US12716548

申请日：2010-03-03

申请人： David J. Wiebe

发明人： David J. Wiebe

IPC分类号： F01D5/18

CPC分类号： F01D5/18

摘要： A cooling system for the fillet of a turbine blade is provided. The blade includes an airfoil transitioning to a platform having a flow path surface. The transition region is defined by a fillet. A cooling passage is formed in the platform and extends about at least a portion of the periphery of the airfoil. The cooling passage is located proximate to the flow path surface and is substantially aligned with at least a portion of the fillet. Coolant is delivered to the passage by a supply hole, which can reduce the temperature in the fillet region. As a result, thermal gradients in the fillet region can be minimized, which can reduce thermal stresses. An exhaust hole extends between the passage and the flow path surface of the platform. Thus, coolant discharged from the exhaust holes enters the flow path of the turbine.

摘要翻译： 提供了一种用于涡轮叶片的圆角的冷却系统。 叶片包括转变到具有流动路径表面的平台的翼型件。 过渡区域由圆角定义。 冷却通道形成在平台中并围绕翼型的周边的至少一部分延伸。 冷却通道位于流路表面附近，并且基本上与圆角的至少一部分对齐。 冷却液通过供应孔输送到通道，这可以降低圆角区域的温度。 因此，可以最小化圆角区域的热梯度，这可以降低热应力。 排气孔在平台的通道和流动通道表面之间延伸。 因此，从排气孔排出的冷却剂进入涡轮机的流路。

公开(公告)号：US20130219921A1

公开(公告)日：2013-08-29

申请号：US13408061

申请日：2012-02-29

申请人： David J. Wiebe ,  Jose L. Rodrigez

发明人： David J. Wiebe ,  Jose L. Rodrigez

IPC分类号： F02C6/08

CPC分类号： F01D9/023 ,  F01D25/12 ,  F02C3/14 ,  F05D2250/314 ,  F23R3/52 ,  F23R3/58

摘要： A cooling system is provided for a transition (420) of a gas turbine engine (410). The cooling system includes a cowling (460) configured to receive an air flow (111) from an outlet of a compressor section of the gas turbine engine (410). The cowling (460) is positioned adjacent to a region of the transition (420) to cool the transition region upon circulation of the air flow within the cowling (460). The cooling system further includes a manifold (121) to directly couple the air flow (111) from the compressor section outlet to an inlet (462) of the cowling (460). The cowling (460) is configured to circulate the air flow (111) within an interior space (426) of the cowling (460) that extends radially outward from an inner diameter (423) of the cowling to an outer diameter (424) of the cowling at an outer surface.

摘要翻译： 为燃气涡轮发动机（410）的过渡（420）提供冷却系统。 冷却系统包括构造成从燃气涡轮发动机（410）的压缩机部分的出口接收空气流（111）的整流罩（460）。 整流罩（460）定位成邻近过渡区域（420），以在气流（460）内的空气流循环时冷却过渡区域。 冷却系统还包括将空气流（111）从压缩机部分出口直接联接到整流罩（460）的入口（462）的歧管（121）。 整流罩（460）构造成使得空气流（111）在整流罩（460）的内部空间（426）内循环，该整流罩（460）从整流罩的内径（423）径向向外延伸到外壳 在外表面的整流罩。

公开(公告)号：US20130081407A1

公开(公告)日：2013-04-04

申请号：US13252348

申请日：2011-10-04

申请人： David J. Wiebe

发明人： David J. Wiebe

IPC分类号： F02C3/04 ,  B23P17/00

CPC分类号： F23R3/46 ,  F01D9/023 ,  F02C3/04 ,  F02C3/14 ,  F05D2230/80 ,  F05D2250/314 ,  Y10T29/49229

摘要： An aero-derivative can annular gas turbine engine having: an aero gas turbine engine core including an aero high pressure compressor (65) interconnected with an aero high pressure turbine (73) by an aero high pressure shaft (142) in a geometric arrangement appropriate for association with an aero annular combustor (84), but with the aero annular combustor (84) and a first row of turbine vanes (38) of the aero high pressure turbine (73) absent; and a can annular combustor assembly (122) assembled with the aero gas turbine engine core and configured to receive compressed air from the aero high pressure compressor (65) and to accelerate and orient combustion gasses directly onto a first row of blades of the aero high pressure turbine (73).

摘要翻译： 一种航空衍生罐环形燃气涡轮发动机，其具有：航空燃气涡轮发动机核心，其包括通过航空高压轴（142）以航空高压涡轮机（142）以几何布置适当的方式与航空高压涡轮机（73）互连的航空高压压缩机（65） 用于与航空环形燃烧器（84）相关联，但是与空气环形燃烧器（84）和空气高压涡轮机（73）的第一排涡轮机叶片（38）不相关; 以及与所述航空燃气涡轮发动机芯组装的并且被配置为从所述航空高压压缩机（65）接收压缩空气并且将燃烧气体直接加速并定向到所述航空高压压缩机（65）的第一排叶片上的罐环形燃烧器组件（122） 压力涡轮机（73）。

公开(公告)号：US08281594B2

公开(公告)日：2012-10-09

申请号：US12555134

申请日：2009-09-08

申请人： David J. Wiebe

发明人： David J. Wiebe

IPC分类号： F02C1/00

CPC分类号： F23D11/36 ,  F23R3/283

摘要： A fuel injector in a combustor apparatus of a gas turbine engine. An outer wall of the injector defines an interior volume in which an intermediate wall is disposed. A first gap is formed between the outer wall and the intermediate wall. The intermediate wall defines an internal volume in which an inner wall is disposed. A second gap is formed between the intermediate wall and the inner wall. The second gap receives cooling fluid that cools the injector. The cooling fluid provides convective cooling to the intermediate wall as it flows within the second gap. The cooling fluid also flows through apertures in the intermediate wall into the first gap where it provides impingement cooling to the outer wall and provides convective cooling to the outer wall. The inner wall defines a passageway that delivers fuel into a liner downstream from a main combustion zone.

摘要翻译： 燃气涡轮发动机的燃烧器装置中的燃料喷射器。 注射器的外壁限定内部容积，其中设置中间壁。 在外壁和中间壁之间形成第一间隙。 中间壁限定了内壁的内部容积。 在中间壁和内壁之间形成第二间隙。 第二间隙接收冷却喷射器的冷却流体。 冷却流体在第二间隙内流动时向中间壁提供对流冷却。 冷却流体还通过中间壁中的孔流入第一间隙，在第一间隙中，其向外壁提供冲击冷却并向外壁提供对流冷却。 内壁限定了将燃料输送到主燃烧区下游的衬套的通道。

公开(公告)号：US20100071377A1

公开(公告)日：2010-03-25

申请号：US12477397

申请日：2009-06-03

申请人： Timothy A. Fox ,  David J. Wiebe

发明人： Timothy A. Fox ,  David J. Wiebe

IPC分类号： F02C7/22 ,  F02C5/02

CPC分类号： F23R3/16 ,  F23R3/283 ,  F23R3/346 ,  F23R2900/00005

摘要： A combustor apparatus for use in a gas turbine engine. The combustor apparatus includes a liner, a flow sleeve, and a fuel injection system. The liner includes an inner volume, wherein a portion of the inner volume defines a main combustion zone. The flow sleeve receives compressed air, is positioned radially outward from the liner, and includes a forward end and an aft end. The fuel injection system is coupled to the flow sleeve and provides fuel into the inner volume of the liner downstream from the main combustion zone. The fuel injection system includes a fuel manifold and a fuel dispensing structure. The fuel manifold is coupled to the flow sleeve and includes a cavity for receiving fuel. The fuel dispensing structure is associated with the cavity and distributes fuel from the cavity to the liner inner volume.

摘要翻译： 一种用于燃气涡轮发动机的燃烧器装置。 燃烧器装置包括衬套，流动套筒和燃料喷射系统。 内衬包括内部容积，其中内部容积的一部分限定主燃烧区。 流动套筒容纳压缩空气，从衬套径向向外定位，并且包括前端和后端。 燃料喷射系统联接到流动套筒并且将燃料提供到主燃烧区下游的衬套的内部容积中。 燃料喷射系统包括燃料歧管和燃料分配结构。 燃料歧管联接到流动套管并且包括用于接收燃料的空腔。 燃料分配结构与空腔相关联，并将燃料从空腔分配到衬里内部容积。