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1. 发明申请

US20100034638A1 Impingement cooling arrangement 审中-公开
标题翻译：冲击冷却装置
公开(公告)号：US20100034638A1
公开(公告)日：2010-02-11
申请号：US11056185
申请日：2005-02-14
申请人： Andrew Chambers , Peter Ireland , David R.H. Gillespie , Geoffrey M. Dailey
发明人： Andrew Chambers , Peter Ireland , David R.H. Gillespie , Geoffrey M. Dailey
IPC分类号： F02C7/12
CPC分类号： F01D25/12 , F01D5/187 , F05D2250/231 , F05D2250/292 , F05D2260/201 , F05D2260/2212 , F05D2260/22141 , Y02T50/673 , Y02T50/675 , Y02T50/676
摘要： An impingement cooling arrangement comprises a projection extending partially across a coolant passage upstream of a jet aperture. An end surface of the projection increases the available surface area for heat exchange with a cross flow whilst a coolant air flow jetted from the jet aperture can transgress a proportion of the air flow passing between the end surface and a junction surface incorporating the jet aperture. A spacing gap B between the end surface and the junction surface avoids localised distortions to the cross flow whilst the projection provides that the coolant air flow projected from the jet aperture mostly passes through a lower turbulence wake downstream of the projection for greater impingement upon a target surface for heat transfer and cooling efficiency. Typically the impingement cooling arrangement is incorporated within turbine blades or vanes of a jet engine.
摘要翻译：冲击冷却装置包括部分地穿过喷射孔上游的冷却剂通道延伸的突出部。突起的端面增加了用于与横流进行热交换的可用表面积，而从喷射孔喷射的冷却剂空气流可以超过在端表面和结合有喷射孔的接合面之间通过的一部分空气流。端表面和接合表面之间的间隔间隙B避免了横向流动的局部变形，同时突出部分提供从喷射孔径突出的冷却剂空气流大部分通过突出部下游的下部湍流尾迹，以便更大程度地撞击目标表面用于传热和冷却效率。通常，冲击冷却装置结合在喷气发动机的涡轮机叶片或叶片内。

2. 发明授权

US06874992B2 Gas turbine engine aerofoil 有权
标题翻译：燃气涡轮发动机机翼
公开(公告)号：US06874992B2
公开(公告)日：2005-04-05
申请号：US10294666
申请日：2002-11-15
申请人： Geoffrey M Dailey
发明人： Geoffrey M Dailey
IPC分类号： F01D5/18 , F01D5/08
CPC分类号： F01D5/187
摘要： An aerofoil blade or vane for a gas turbine engine comprises a body member having an inner end for mounting the blade on a shaft and an outer or tip end. A plurality of cooling passages are formed within the blade, the cooling passages comprising a plurality of inlet passages along which cooling air flows from the base towards the tip region of the blade and a plurality of return passages along which cooling air flows from the tip towards the base region of the blade. At least some of the passages are connected by a common chamber located within the tip region of the blade.
摘要翻译：用于燃气涡轮发动机的机翼叶片或叶片包括主体构件，其具有用于将叶片安装在轴上的内端和外端或尖端。多个冷却通道形成在叶片内，冷却通道包括多个入口通道，冷却空气从基座朝向叶片的顶端区域流过冷却空气，多个返回通道沿着冷却空气从顶端流向叶片的基部区域。至少一些通道通过位于叶片的尖端区域内的公共室连接。

3. 发明授权

US06564557B1 Cooling arrangement 失效
标题翻译：冷却布置
公开(公告)号：US06564557B1
公开(公告)日：2003-05-20
申请号：US09910060
申请日：2001-07-23
申请人： Geoffrey M Dailey
发明人： Geoffrey M Dailey
IPC分类号： F02C712
CPC分类号： F01D25/12 , F02C7/18
摘要： Cooling air enters a cooling path 56 from a supply at 58. The path follows a serpentine path consecutively through a number of components to be cooled, including guide vanes 38, 40, blades 30, a shroud 54 and a shaft 18. A valve 68 is provided at the downstream end of the path 56. The setting of the valve 56 modulates the flow and pressure of air along the path 56, without wholly preventing it.
摘要翻译：冷却空气从58处的供应部分进入冷却路径56.路径沿着蛇形路径连续地穿过要冷却的多个部件，包括导向叶片38,40，叶片30，护罩54和轴18.阀68 设置在路径56的下游端。阀56的设定调节沿着路径56的空气的流量和压力，而不完全阻止它。

4. 发明授权

US06554570B2 Turbine blade support assembly and a turbine assembly 失效
公开(公告)号：US06554570B2
公开(公告)日：2003-04-29
申请号：US09923338
申请日：2001-08-08
申请人： Geoffrey M Dailey
发明人： Geoffrey M Dailey
IPC分类号： B64C1100
CPC分类号： F01D5/182 , F01D5/18 , F01D5/186 , F01D5/187 , F05D2260/201 , F05D2260/202 , F05D2260/203 , F05D2260/2212
摘要： A turbine assembly (35) for a gas turbine engine (10) comprises a rotatable support arrangement (38) which comprises means for mounting thereon a plurality of turbine blades (36). The turbine assembly (35) defines flow path means (43) for a flow of cooling fluid therethrough. The flow path means (43) is connectable to a supply of relatively cold cooling fluid. The flow path means 43 is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means (43) substantially wholly by the centrifugal force generated the rotation of the turbine assembly (35) in operation. Relatively hot cooling fluid is displaced by the relatively cold cooling fluid radially inwardly through the flow path means (43).

5. 发明授权

US06506020B2 Blade platform cooling 有权
标题翻译：刀片平台冷却
公开(公告)号：US06506020B2
公开(公告)日：2003-01-14
申请号：US09901075
申请日：2001-07-10
申请人： Geoffrey M Dailey
发明人： Geoffrey M Dailey
IPC分类号： F01D518
CPC分类号： F01D5/18
摘要： A turbine assembly for a gas turbine engine includes a plurality of turbine blades (32) mounted on a rotatable support means in the form of a turbine disc so as to extend radially therefrom. The turbine blades include circumferentially extending blade platforms (40) spaced from the turbine disc and means are provided for allowing the passage of air between an internal region of the blades (32) and a space located between the blade platforms (40) and the turbine disc. The air may flow out of and back into the same turbine blade, or may flow into an adjacent blade. This flow of air results in the cooling of the blade platforms (40).
摘要翻译：用于燃气涡轮发动机的涡轮机组件包括多个涡轮机叶片（32），其安装在涡轮盘形式的可旋转支撑装置上，以便从其径向延伸。涡轮叶片包括与涡轮盘间隔开的周向延伸的叶片平台（40），并且提供装置用于允许空气在叶片（32）的内部区域和位于叶片平台（40）和涡轮机光盘。空气可能流出并返回到相同的涡轮叶片中，或者可以流入相邻的叶片。这种空气流导致叶片平台（40）的冷却。

6. 发明授权

US06837683B2 Gas turbine engine aerofoil 有权
标题翻译：燃气涡轮发动机机翼
公开(公告)号：US06837683B2
公开(公告)日：2005-01-04
申请号：US10291408
申请日：2002-11-12
申请人： Geoffrey M Dailey
发明人： Geoffrey M Dailey
IPC分类号： F01D5/18 , F01D5/08
CPC分类号： F01D5/187 , F05D2260/201 , F05D2260/202
摘要： A gas turbine engine blade or vane comprises inner linked chambers. A chamber adjacent the leading edge is provided with an inlet for receiving cooling fluid and a chamber adjacent the trailing edge is provided with an outlet for exhausting cooling fluid. The chambers are arranged in series from the leading edge to the trailing edge so as to direct cooling fluid within the aerofoil blade or vane from the leading edge region to the trailing edge region.
摘要翻译：燃气涡轮发动机叶片或叶片包括内连接室。邻近前缘的腔室设置有用于接收冷却流体的入口，并且与后缘相邻的腔室设置有用于排出冷却流体的出口。这些室从前缘到后缘串联布置，以便将机翼叶片或叶片内的冷却流体从前缘区域引导到后缘区域。

7. 发明授权

US06334297B1 Combuster arrangement 有权
标题翻译：燃烧器安排
公开(公告)号：US06334297B1
公开(公告)日：2002-01-01
申请号：US09624331
申请日：2000-07-24
申请人： Geoffrey M Dailey , Anthony Pidcock , Desmond Close
发明人： Geoffrey M Dailey , Anthony Pidcock , Desmond Close
IPC分类号： F02C718
CPC分类号： F23R3/04 , F23R3/346 , F23R2900/03041 , Y02T50/675
摘要： A combustor arrangement (8) for a gas turbine engine (3) comprising a combustion chamber, fuel nozzles(52a,52b), and a bled diffuser (7) located upstream of said combustion chamber to, in use, direct an airflow from an upstream compressor (6) into the combustor (8). The fuel nozzles (52a,52b) arranged in use to supply fuel into the combustion chamber where it is mixed and combusted with the airflow from the compressor (6). The bled diffuser (7) adapted to bleed off a portion of said airflow from a main airflow into the combustion chamber. At least one bleed duct (46) is connected to the bled diffuser (7) to return and direct the air bled from the diffuser (7) to a main gas flow through the engine (3) at a location downstream of the fuel supply means (52a,52b). The bled air from the diffuser (7) preferably providing cooling of a part of the gas turbine engine (3), for example part of the turbine (10) or combustor outlet vane (58), downstream of the combustor (8).
摘要翻译：一种用于燃气涡轮发动机（3）的燃烧器装置（8），其包括燃烧室，燃料喷嘴（52a，52b）和位于所述燃烧室上游的流体扩散器（7），在使用中，引导来自上游压缩机（6）进入燃烧器（8）。燃料喷嘴（52a，52b），其布置成用于将燃料供应到燃烧室中，在该燃烧室中与来自压缩机（6）的气流混合并燃烧。流血扩散器（7）适于将一部分所述气流从主气流排出到燃烧室中。至少一个排放管道（46）连接到排放的扩散器（7），以在燃料供应装置的下游位置返回并引导从扩散器（7）排出的空气通过发动机（3）的主气体流（52a，52b）。来自扩散器（7）的排出的空气优选地提供燃气涡轮发动机（3）的一部分的冷却，例如在燃烧器（8）下游的涡轮机（10）或燃烧器出口叶片（58）的一部分。

8. 发明授权

US06264428B1 Cooled aerofoil for a gas turbine engine 有权
标题翻译：用于燃气涡轮发动机的冷却机翼
公开(公告)号：US06264428B1
公开(公告)日：2001-07-24
申请号：US09480691
申请日：2000-01-11
申请人： Geoffrey M Dailey , Peter A Evans , Richard A B McCall
发明人： Geoffrey M Dailey , Peter A Evans , Richard A B McCall
IPC分类号： F01D518
CPC分类号： F01D5/187 , Y02T50/673 , Y02T50/676
摘要： A hollow cooled aerofoil blade (10) is provided with a central, lengthways extending plenum (16) and has pressure and suction flanks (20,15). The suction flank (15) has lengthways extending passages (21) therein that direct cooling air from the root region (13) into the central plenum (16). From the central plenum, the cooling air flows through a first set of apertures (27) into lengthways extending passages (28) in the pressure flank (20). The cooling air then flows through a second set of apertures (30) in the pressure flank to provide film cooling of the pressure flank (20) external surface.
摘要翻译：中空的冷却机翼叶片（10）设置有中心的纵向延伸的气室（16）并具有压力和吸力侧面（20,15）。吸力侧面（15）在其中具有纵向延伸的通道（21），其将冷却空气从根部区域（13）引导到中央增压室（16）中。冷却空气从中央集气室经过第一组孔（27）流入压力侧面（20）中的纵向延伸通道（28）。冷却空气然后流过压力侧面中的第二组孔（30），以提供压力侧面（20）外表面的膜冷却。

9. 发明授权

US08973371B2 Gas turbine engine with turbine cooling arrangement 有权
标题翻译：带涡轮冷却装置的燃气轮机
公开(公告)号：US08973371B2
公开(公告)日：2015-03-10
申请号：US13224894
申请日：2011-09-02
申请人： Jonathan M King , Crispin D. Bolgar , Guy D. Snowsill , Michael J. Sheath , Geoffrey M Dailey
发明人： Jonathan M King , Crispin D. Bolgar , Guy D. Snowsill , Michael J. Sheath , Geoffrey M Dailey
IPC分类号： F02C6/08 , F02C3/13 , F02C7/18 , F01D5/08 , F01D11/00 , F01D11/04 , F01D11/06
CPC分类号： F02C6/08 , F01D5/081 , F01D5/082 , F01D11/001 , F01D11/04 , F01D11/06 , F02C3/13 , F02C7/18
摘要： A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to form a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.
摘要翻译：提供一种燃气涡轮发动机，其包括涡轮部分冷却系统和冷却燃气涡轮发动机的涡轮部分的方法。燃气涡轮发动机包括流动系列，压缩机部分，燃烧器和涡轮部分，发动机还包括涡轮部分冷却系统。涡轮部分冷却系统包括第一压缩空气排放装置和第二压缩空气排放装置。第一压缩排气装置从压缩机部分的高压级排出第一流的压缩空气。压缩空气的第一流动绕过燃烧器并到达涡轮部分，以在相邻转子盘的上游的一排定子叶片上形成密封和/或冷却流。第二压缩空气排出装置从压缩机部分的一个或多个较低压力级渗出第二压缩空气流。压缩空气的第二流动绕过燃烧器并到达涡轮部分以形成冷却流。冷却流的第一部分被引导到转子盘的前表面，并且冷却流的第二部分被引导到转子盘的后表面。

10. 发明申请

US20120060507A1 GAS TURBINE ENGINE 有权
标题翻译：气体涡轮发动机
公开(公告)号：US20120060507A1
公开(公告)日：2012-03-15
申请号：US13224894
申请日：2011-09-02
申请人： Jonathan M. KING , Crispin D. BOLGAR , Guy D. SNOWSILL , Michael J. SHEATH , Geoffrey M. DAILEY
发明人： Jonathan M. KING , Crispin D. BOLGAR , Guy D. SNOWSILL , Michael J. SHEATH , Geoffrey M. DAILEY
IPC分类号： F02C7/18 , F02C3/04
CPC分类号： F02C6/08 , F01D5/081 , F01D5/082 , F01D11/001 , F01D11/04 , F01D11/06 , F02C3/13 , F02C7/18
摘要： A gas turbine engine comprising a turbine section cooling system and a method of cooling a turbine section of a gas turbine engine is provided. The gas turbine engine comprises in flow series a compressor section, a combustor, and a turbine section, the engine further comprising a turbine section cooling system. The turbine section cooling system including a first compressed air bleed arrangement and a second compressed air bleed arrangement. The first compressed air bleed arrangement bleeds a first flow of compressed air from a high pressure stage of the compressor section. The first flow of compressed air bypasses the combustor and arrives at the turbine section to form a sealing and/or cooling flow at a row of stator vanes upstream of an adjacent rotor disc. The second compressed air bleed arrangement bleeds a second flow of compressed air from one or more lower pressure stages of the compressor section. The second flow of compressed air bypasses the combustor and arrives at the turbine section to fowl a cooling flow. A first portion of the cooling flow is routed to a front face of the rotor disc and a second portion of the cooling flow is routed to a rear face of the rotor disc.
摘要翻译：提供一种燃气涡轮发动机，其包括涡轮部分冷却系统和冷却燃气涡轮发动机的涡轮部分的方法。燃气涡轮发动机包括流动系列，压缩机部分，燃烧器和涡轮部分，发动机还包括涡轮部分冷却系统。涡轮部分冷却系统包括第一压缩空气排放装置和第二压缩空气排放装置。第一压缩排气装置从压缩机部分的高压级排出第一流的压缩空气。压缩空气的第一流动绕过燃烧器并到达涡轮部分以在相邻转子盘的上游的一排定子叶片处形成密封和/或冷却流。第二压缩空气排出装置从压缩机部分的一个或多个较低压力级渗出第二压缩空气流。压缩空气的第二流动绕过燃烧器并到达涡轮部分以产生冷却流。冷却流的第一部分被引导到转子盘的前表面，并且冷却流的第二部分被引导到转子盘的后表面。

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