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1. 发明申请

US20130219921A1 MID-SECTION OF A CAN-ANNULAR GAS TURBINE ENGINE WITH A COOLING SYSTEM FOR THE TRANSITION 有权
标题翻译：具有用于过渡的冷却系统的CAN-ANNULAR气体涡轮发动机的中间部分
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。

2. 发明申请

US20150000283A1 COMBUSTOR APPARATUS IN A GAS TURBINE ENGINE 有权
标题翻译：燃气涡轮发动机中的燃烧器装置
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。

3. 发明授权

US08668454B2 Turbine airfoil fillet cooling system 失效
标题翻译：涡轮机翼角冷却系统
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。

4. 发明申请

US20130081407A1 AERO-DERIVATIVE GAS TURBINE ENGINE WITH AN ADVANCED TRANSITION DUCT COMBUSTION ASSEMBLY 审中-公开
标题翻译：具有先进过渡式燃煤组件的AERO-DERIVATIVE GAS TURBINE发动机
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）。

5. 发明授权

US08281594B2 Fuel injector for use in a gas turbine engine 有权
标题翻译：用于燃气涡轮发动机的燃油喷射器
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。

6. 发明申请

US20100071377A1 Combustor Apparatus for Use in a Gas Turbine Engine 审中-公开
标题翻译：用于燃气轮机发动机的燃烧器装置
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。

7. 发明授权

US5314304A Abradeable labyrinth stator seal 失效
标题翻译：耐磨迷宫式定子密封
公开(公告)号：US5314304A
公开(公告)日：1994-05-24
申请号：US745630
申请日：1991-08-15
申请人： David J. Wiebe
发明人： David J. Wiebe
IPC分类号： F01D11/12 , F01D11/08
CPC分类号： F01D11/127 , F01D11/122 , Y10T428/1234 , Y10T428/12611
摘要： In a gas turbine engine having a laybrinth seal between an annular rotor and stator therein, in which the stator is surmounted by a honeycomb structure and the rotor has a knife edge which is mounted to rotate in close annular proximity with said honeycomb structure, an improvement is provided wherein such labyrinth seal has a layer of abradeable coating atop the honeycomb structure for the rotor knife edge to rotate proximate thereto and to rub in, without substantially damaging the knife edge nor the honeycomb structure. In one embodiment the layer of abradeable coating is mounted on a metalic foil which is mounted in turn, atop the honeycomb structure. The abradeable coating thus provided, is more yieldable and less damaging to a rotor knife edge than is the honeycomb structure of the prior art, to better preserve rotor and stator and thus the labryrinth seal. Because of such protection, the rotor knife edge can be thinner and of lighter weight and the honeycomb structure can be made of larger cell sizes, again resulting in weight savings for each lab seal, which can have one or a plurality of rotor (knife edge)-stator pairs. Further, the abradable coating seals the top of the honeycomb structure and thus blocks air flow into the honeycomb cells beneath and behind the knife edge and thus reduces losses in seal efficiency.
摘要翻译：在具有在其中的环形转子和定子之间具有底孔密封的燃气涡轮发动机中，其中定子被蜂窝结构覆盖，并且转子具有安装成与所述蜂窝结构紧密环接地旋转的刀刃，改进其中这种迷宫式密封件具有在蜂窝状结构顶部的可磨蚀涂层，用于转子刀刃在其附近旋转并摩擦，而基本上不损坏刀刃和蜂窝结构。在一个实施方案中，可磨蚀涂层是安装在依次安装在蜂窝结构顶上的金属箔上的。如此提供的可磨损涂层比现有技术的蜂窝结构更可屈服并且对转子刀刃的损伤较小，以更好地保持转子和定子，从而更好地保持了实验室密封。由于这样的保护，转子刀刃可以更薄和更轻的重量，并且蜂窝结构可以由更大的电池尺寸制成，这再次导致每个实验室密封件的重量节省，其可以具有一个或多个转子（刀刃）-stator对。此外，可磨损涂层密封蜂窝结构体的顶部，从而阻挡在刀刃下方和后方的进入蜂窝状电池的空气流，从而减少密封效率的损失。

8. 发明授权

US09399926B2 Belly band seal with circumferential spacer 有权
标题翻译：带周边垫圈的带状密封圈
公开(公告)号：US09399926B2
公开(公告)日：2016-07-26
申请号：US13974147
申请日：2013-08-23
申请人： David J. Wiebe
发明人： David J. Wiebe
IPC分类号： F01D11/08 , F01D5/06 , F01D11/00 , F02C7/28
CPC分类号： F01D11/08 , F01D5/06 , F01D5/066 , F01D11/003 , F01D11/005 , F01D11/008 , F02C7/28
摘要： A circumferentially extending sealing band is located within sealing band receiving slots formed in adjacent turbine engine disks. The sealing band includes a plurality of seal strips forming overlap joints defined by overlapping end portions, each formed by a tongue portion extending from a seal face of one seal strip past a seal face of the adjacent seal strip, along a radially inward facing side of the adjacent seal strip. A joint gap is defined within at least one overlap joint between the seal face of the one seal strip and the seal face of the adjacent strip, and a spacer is affixed to the one seal strip and is located at a position within the joint gap between the seal faces to limit circumferential movement of the seal faces toward each other.
摘要翻译：周向延伸的密封带位于形成在相邻的涡轮引擎盘中的密封带接收槽内。密封带包括形成由重叠端部限定的重叠接头的多个密封条，每个密封条由从一个密封条的密封面延伸穿过相邻密封条的密封面的舌部形成，该舌部沿着径向向内的侧面相邻密封条。在一个密封条的密封面和相邻条的密封面之间的至少一个重叠接合部内限定了接合间隙，并且间隔件固定到一个密封条上，并且位于第一密封条之间的接合间隙内的位置密封面面对以限制密封面朝向彼此的周向移动。

9. 发明授权

US09267689B2 Combustor apparatus in a gas turbine engine 有权
标题翻译：燃气涡轮发动机中的燃烧器装置
公开(公告)号：US09267689B2
公开(公告)日：2016-02-23
申请号：US13783857
申请日：2013-03-04
申请人： David J. Wiebe , Stephen A. Ramier , Robert H. Bartley , Daniel W. Garan , Erick J. Deane , Mark L. Adamson , Yadollah Naghian
发明人： David J. Wiebe , Stephen A. Ramier , Robert H. Bartley , Daniel W. Garan , Erick J. Deane , Mark L. Adamson , Yadollah Naghian
IPC分类号： F23R3/28 , F02C7/22 , F16L13/007 , F23R3/34 , F16L41/08
CPC分类号： F23R3/28 , F02C7/22 , F02C7/222 , F05D2300/50211 , F05D2300/50212 , F05D2300/506 , F16L13/007 , F16L15/006 , F16L19/061 , F16L25/0072 , F16L41/084 , F16L41/086 , F23K2301/20 , F23R3/283 , F23R3/286 , F23R3/34 , F23R2900/00005
摘要： A combustor apparatus defines a combustion zone where air and fuel are burned to create high temperature combustion products. The combustor apparatus includes an outer wall, coupling structure on the outer wall adjacent to a fuel inlet opening thereof, a fuel injection system, a fuel feed assembly, and a fitting member. The fuel injection system provides fuel to be burned in the combustion zone. The fuel supply structure includes a threaded inner surface formed from a first material. The fuel feed assembly includes a fuel feed pipe that extends through the fuel inlet opening in the outer wall and has an outlet portion formed from the first material and that is threadedly engaged with the fuel supply structure, and an inlet portion affixed to the outlet portion and formed from a second material. The fitting member secures the fuel feed assembly relative to the outer wall.
摘要翻译：燃烧器装置限定了燃烧区域，其中空气和燃料被燃烧以产生高温燃烧产物。燃烧器装置包括外壁，与其燃料入口相邻的外壁上的联接结构，燃料喷射系统，燃料供给组件和配件构件。燃料喷射系统提供在燃烧区域燃烧的燃料。燃料供给结构包括由第一材料形成的螺纹内表面。所述燃料供给组件包括燃料供给管，所述燃料供给管延伸穿过所述外壁中的所述燃料入口，并且具有由所述第一材料形成并与所述燃料供给结构螺纹接合的出口部，以及固定到所述出口部的入口部并由第二材料形成。装配构件相对于外壁固定燃料供给组件。

10. 发明申请

US20150107264A1 STRUCTURAL MOUNTING ARRANGEMENT FOR GAS TURBINE ENGINE COMBUSTION GAS DUCT 有权
标题翻译：气体涡轮发动机燃烧气体结构安装布置
公开(公告)号：US20150107264A1
公开(公告)日：2015-04-23
申请号：US14059521
申请日：2013-10-22
申请人： David J. Wiebe , Richard C. Charron , Jay A. Morrison
发明人： David J. Wiebe , Richard C. Charron , Jay A. Morrison
IPC分类号： F23R3/60 , F01D9/02
CPC分类号： F23R3/60 , F01D9/023 , F05D2260/30 , F23R3/46
摘要： A gas turbine engine ducting arrangement (10), including: an annular chamber (14) configured to receive a plurality of discrete flows of combustion gases originating in respective can combustors and to deliver the discrete flows to a turbine inlet annulus, wherein the annular chamber includes an inner diameter (52) and an outer diameter (60); an outer diameter mounting arrangement (34) configured to permit relative radial movement and to prevent relative axial and circumferential movement between the outer diameter and a turbine vane carrier (20); and an inner diameter mounting arrangement (36) including a bracket (64) secured to the turbine vane carrier, wherein the bracket is configured to permit the inner diameter to move radially with the outer diameter and prevent axial deflection of the inner diameter with respect to the outer diameter.
摘要翻译：一种燃气涡轮发动机管道装置（10），包括：环形室（14），其构造成容纳来自相应罐式燃烧器的多个不连续的燃烧气体流并且将离散流输送到涡轮机入口环，其中环形室包括内径（52）和外径（60）; 外径安装装置（34）被构造成允许相对径向移动并且防止外径与涡轮叶片载体（20）之间的相对轴向和周向移动; 以及内径安装装置（36），其包括固定到所述涡轮机叶片承载件的支架（64），其中所述支架构造成允许所述内径径向与所述外径径向移动，并且防止所述内径相对于外径。

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