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公开(公告)号：US20100303635A1
公开(公告)日：2010-12-02
申请号：US12787758
申请日：2010-05-26
Applicant: Roderick M. TOWNES , Ian Tibbott , Edwin Dane , Caner H. Helvaci
Inventor： Roderick M. TOWNES , Ian Tibbott , Edwin Dane , Caner H. Helvaci
IPC: F01D5/18
CPC classification number: F01D5/187 , F05D2250/11 , F05D2250/70 , F05D2260/2212 , F05D2260/22141
Abstract: Providing cooling within hollow blades such as high pressure turbine blades in a gas turbine engine is important to maintain these components within operational margins for the materials from which they are formed. Traditionally, coolant flows in hollow passages have been used along with impingement apertures towards a leading passage for cooling effectiveness. It is known that opposed undulations or ribs can create rotational vortices within the passage. By shaping shaped portions between the opposed undulations and possibly providing undulations upon these shaped portions themselves it is possible to generate stronger more powerful vortices within the passage. These vortices are coupled with the impingement orifices to create proportionally greater impingement jet flow and pressure and therefore cooling effectiveness within the leading passage.
Abstract translation: 在诸如燃气轮机发动机中的高压涡轮机叶片的中空叶片内提供冷却对于将这些部件保持在其形成的材料的操作边缘内是重要的。 传统上，在中空通道中的冷却剂流已经与冲击孔一起使用，朝向前导通道以降低冷却效果。 已知相对的起伏或肋可以在通道内产生旋转涡流。 通过在相对的波纹之间成形成形部分并且可能在这些成形部分本身上提供起伏，可以在通道内产生更强大的更强大的涡流。 这些涡流与冲击孔结合，以产生比例地更大的冲击射流和压力，从而在前导通道内产生冷却效果。
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公开(公告)号：US20070253815A1
公开(公告)日：2007-11-01
申请号：US11207892
申请日：2005-08-22
Applicant: Michiel Kopmels , Ian Tibbott , Edwin Dane , Mark Mitchell
Inventor： Michiel Kopmels , Ian Tibbott , Edwin Dane , Mark Mitchell
IPC: F01D5/18
Abstract: An internal fluid cooling system for a gas turbine aerofoil (2) comprises a plurality of multi-pass cooling arrangements each of which consists of a serpentine passage (50,52,54 & 48,56,58) in the interior of the aerofoil (2). The cooling fluid—in particular air—is supplied to an inlet end (22,24) of each passage (50,52,54 & 48,56,58) and exhausted through a multiplicity of discharge holes (60,62,64,68) to provide tip, leading edge, trailing edge and surface film cooling. The inlet end (22) of a first serpentine passage (50,52,54) is positioned close to the leading edge (26) and flows rearwards while the inlet end (24) of the second serpentine passage (48,56,58) is positioned close to the trailing edge (28) and flows forwards. These serpentine passages are disposed side-by-side, one adjacent the pressure surface (14) and the other adjacent the suction surface (16) on opposite sides of a main load carrying member (30) which comprises a major part of the internal structure (30,36,38,40,42,44) of the aerofoil (2).
Abstract translation: 用于燃气轮机机翼（2）的内部流体冷却系统包括多个多通道冷却装置，每个多通道冷却装置由机翼内部的蛇形通道（50,52,54和48,56,58） 2）。 冷却流体 - 特别是空气 - 被供应到每个通道（50,52,54和48,56,58）的入口端（22,24），并通过多个排放孔（60,62,64， 68）提供尖端，前缘，后缘和表面膜冷却。 第一蛇形通道（50,52,54）的入口端（22）定位成靠近前缘（26）并向后流动，同时第二蛇形通道（48,56,58）的入口端（24） 定位成靠近后缘（28）并向前流动。 这些蛇形通道并排设置，一个邻近压力表面（14），另一个邻近主负载承载构件（30）的相对侧的吸附表面（16），该主负载承载构件（30）包括内部结构的主要部分 （30,36,38,40,42,44）的机翼（2）。
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公开(公告)号：US08523523B2
公开(公告)日：2013-09-03
申请号：US12787758
申请日：2010-05-26
Applicant: Roderick M. Townes , Ian Tibbott , Edwin Dane , Caner H. Helvaci
Inventor： Roderick M. Townes , Ian Tibbott , Edwin Dane , Caner H. Helvaci
CPC classification number: F01D5/187 , F05D2250/11 , F05D2250/70 , F05D2260/2212 , F05D2260/22141
Abstract: Providing cooling within hollow blades such as high pressure turbine blades in a gas turbine engine is important to maintain these components within operational margins for the materials from which they are formed. Traditionally, coolant flows in hollow passages have been used along with impingement apertures towards a leading passage for cooling effectiveness. It is known that opposed undulations or ribs can create rotational vortices within the passage. By shaping shaped portions between the opposed undulations and possibly providing undulations upon these shaped portions themselves it is possible to generate stronger more powerful vortices within the passage. These vortices are coupled with the impingement orifices to create proportionally greater impingement jet flow and pressure and therefore cooling effectiveness within the leading passage.
Abstract translation: 在诸如燃气轮机发动机中的高压涡轮机叶片的中空叶片内提供冷却对于将这些部件保持在其形成的材料的操作边缘内是重要的。 传统上，在中空通道中的冷却剂流已经与冲击孔一起使用，朝向前导通道以降低冷却效果。 已知相对的起伏或肋可以在通道内产生旋转涡流。 通过在相对的波纹之间成形成形部分并且可能在这些成形部分本身上提供起伏，可以在通道内产生更强大的更强大的涡流。 这些涡流与冲击孔结合，以产生比例地更大的冲击射流和压力，从而在前导通道内产生冷却效果。
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公开(公告)号：US08133032B2
公开(公告)日：2012-03-13
申请号：US12327030
申请日：2008-12-03
Applicant: Ian Tibbott , Edwin Dane , Dougal Richard Jackson
Inventor： Ian Tibbott , Edwin Dane , Dougal Richard Jackson
CPC classification number: F01D5/20 , F01D11/08 , Y02T50/673
Abstract: A rotor blade has a tip with an outer face including at least two channels which each extend to an outlet in the vicinity of the trailing edge. Accordingly, gas leakage around the tip must cross at least three walls, at least in the vicinity of the uncovered turning region near the trailing edge of the blade. Leakage gas entering the channels will tend to create a vortex and pass along the channel to the outlet.
Abstract translation: 转子叶片具有带有外表面的末端，该外表面包括至少两个通道，每个通道延伸到后缘附近的出口。 因此，尖端附近的气体泄漏必须至少跨越至少三个壁，至少在叶片后缘附近的未覆盖的转动区域附近。 进入通道的泄漏气体将倾向于产生涡流并沿着通道传递到出口。
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公开(公告)号：US07850428B2
公开(公告)日：2010-12-14
申请号：US11707997
申请日：2007-02-20
Applicant: Ian Tibbott , Edwin Dane
Inventor： Ian Tibbott , Edwin Dane
CPC classification number: F01D5/186 , F05D2240/12 , F05D2240/122 , F05D2240/304 , F05D2250/16 , F05D2250/71 , F05D2260/221
Abstract: An aerofoil 20 for a gas turbine engine includes a root portion 22, a tip portion 24 located radially outwardly of the root portion 22, leading and trailing edges 26, 28 extending between the root portion 22 and the tip portion 24 and an internal cooling passage 34. The aerofoil 20 includes a plurality of cooling fluid discharge apertures 36 extending between the root portion 22 and the tip portion 24 in a trailing edge region 28a to discharge cooling fluid from the internal cooling passage 34 to an outer surface 31 of the aerofoil in the trailing edge region 28a and thereby provide a cooling film in the trailing edge region 28a. The cooling fluid discharge apertures 36 are arranged so that the flow rate of the cooling fluid discharged from the internal cooling passage 34 to the outer surface trailing edge region 28a varies between the root portion 22 and the tip portion 24.
Abstract translation: 用于燃气涡轮发动机的机翼20包括根部22，位于根部22的径向外侧的前端部24，在根部22和顶端部24之间延伸的前缘26和后缘26以及内部冷却通道 机翼20包括在后缘区域28a中在根部22和尖端部分24之间延伸的多个冷却流体排放孔36，以将冷却流体从内部冷却通道34排出到机翼的外表面31 后缘区域28a，从而在后缘区域28a中提供冷却膜。 冷却流体排出孔36被布置成使得从内部冷却通道34排出到外表面后缘区域28a的冷却流体的流量在根部22和尖端部24之间变化。
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Patent Agency Ranking
公开(公告)号：US20090162200A1
公开(公告)日：2009-06-25
申请号：US12327030
申请日：2008-12-03
Applicant: Ian Tibbott , Edwin Dane , Dougal Richard Jackson
Inventor： Ian Tibbott , Edwin Dane , Dougal Richard Jackson
IPC: B64C11/24
CPC classification number: F01D5/20 , F01D11/08 , Y02T50/673
Abstract: A rotor blade has a tip with an outer face including at least two channels which each extend to an outlet in the vicinity of the trailing edge. Accordingly, gas leakage around the tip must cross at least three walls, at least in the vicinity of the uncovered turning region near the trailing edge of the blade. Leakage gas entering the channels will tend to create a vortex and pass along the channel to the outlet.
Abstract translation: 转子叶片具有带有外表面的末端，该外表面包括至少两个通道，每个通道延伸到后缘附近的出口。 因此，尖端附近的气体泄漏必须至少跨越至少三个壁，至少在叶片后缘附近的未覆盖的转动区域附近。 进入通道的泄漏气体将倾向于产生涡流并沿着通道传递到出口。
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公开(公告)号：US08419366B2
公开(公告)日：2013-04-16
申请号：US12458412
申请日：2009-07-10
Applicant: Roderick M. Townes , Ian Tibbott , Edwin Dane
Inventor： Roderick M. Townes , Ian Tibbott , Edwin Dane
IPC: F01D5/18
CPC classification number: F01D5/186 , F01D5/187 , Y02T50/676
Abstract: Cooling arrangements for blades, and in particular turbine blades utilizing gas turbine engines include impingement apertures with impingement jets, which improve cooling efficiency. By providing a leading passage, which is divided at least into a lower section and an upper section, the lower section can have a wall, which is solid for structural integrity while an upper section has impingement apertures for greater cooling efficiency.
Abstract translation: 叶片，特别是利用燃气涡轮发动机的涡轮叶片的冷却装置包括具有冲击射流的冲击孔，这提高了冷却效率。 通过提供至少分为下部和上部的引导通道，下部可具有壁，该壁为结构完整性而牢固，而上部具有用于更高冷却效率的冲击孔。
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公开(公告)号：US07811058B2
公开(公告)日：2010-10-12
申请号：US11589233
申请日：2006-10-30
Applicant: Ian Tibbott , Edwin Dane
Inventor： Ian Tibbott , Edwin Dane
CPC classification number: F01D5/081 , F01D5/22 , F01D11/008 , F05D2240/81 , F05D2250/314 , Y10S416/50
Abstract: Cooling with regard to high-pressure turbine platforms is important in order to maintain gas turbine engine efficiency. Cottage Roof dampers located below junction gaps between adjacent platforms have been used but tend to present spent coolant flow at a high angle relative to hot gas flows about the aerofoil blades. The present arrangement has the junction gap angled such that the emergent coolant flow remains adjacent to the suction side to create a coolant film lingering above that suction side of the platform.
Abstract translation: 关于高压涡轮平台的冷却对于维持燃气轮机的发动机效率很重要。 小屋位于相邻平台之间的连接间隙以下的屋顶阻尼器已被使用，但倾向于相对于围绕翼型叶片的热气流以高角度呈现废冷却剂流。 本布置具有接合间隙成角度，使得紧急冷却剂流保持邻近吸入侧，以产生在平台的吸力侧上方延伸的冷却剂膜。
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公开(公告)号：US20080273988A1
公开(公告)日：2008-11-06
申请号：US11707997
申请日：2007-02-20
Applicant: Ian Tibbott , Edwin Dane
Inventor： Ian Tibbott , Edwin Dane
IPC: F01D5/18
CPC classification number: F01D5/186 , F05D2240/12 , F05D2240/122 , F05D2240/304 , F05D2250/16 , F05D2250/71 , F05D2260/221
Abstract: An aerofoil 20 for a gas turbine engine includes a root portion 22, a tip portion 24 located radially outwardly of the root portion 22, leading and trailing edges 26, 28 extending between the root portion 22 and the tip portion 24 and an internal cooling passage 34. The aerofoil 20 includes a plurality of cooling fluid discharge apertures 36 extending between the root portion 22 and the tip portion 24 in a trailing edge region 28a to discharge cooling fluid from the internal cooling passage 34 to an outer surface 31 of the aerofoil in the trailing edge region 28a and thereby provide a cooling film in the trailing edge region 28a. The cooling fluid discharge apertures 36 are arranged so that the flow rate of the cooling fluid discharged from the internal cooling passage 34 to the outer surface trailing edge region 28a varies between the root portion 22 and the tip portion 24.
Abstract translation: 用于燃气涡轮发动机的机翼20包括根部22，位于根部22的径向外侧的前端部24，在根部22和顶端部24之间延伸的前缘26和后缘26以及内部冷却通道 机翼20包括在后缘区域28a中在根部22和末端部分24之间延伸的多个冷却流体排出孔36，以将冷却流体从内部冷却通道34排出到机翼的外表面31 在后缘区域28a中，从而在后缘区域28a中提供冷却膜。 冷却流体排出孔36被布置成使得从内部冷却通道34排出到外表面后缘区域28a的冷却流体的流量在根部22和末端部分24之间变化。
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公开(公告)号：US07442008B2
公开(公告)日：2008-10-28
申请号：US11207892
申请日：2005-08-22
Applicant: Michiel Kopmels , Ian Tibbott , Edwin Dane , Timothy M Mitchell
Inventor： Michiel Kopmels , Ian Tibbott , Edwin Dane , Timothy M Mitchell
IPC: F01D5/18
Abstract: An internal fluid cooling system for a gas turbine aerofoil comprises a plurality of multi-pass cooling arrangements each of which consists of a serpentine passage in the interior of the aerofoil. The cooling fluid—in particular air—is supplied to an inlet end of each passage and exhausted through a multiplicity of discharge holes to provide tip, leading edge, trailing edge and surface film cooling. The inlet end of a first serpentine passage is positioned close to the leading edge and flows rearwards while the inlet end of the second serpentine passage is positioned close to the trailing edge and flows forwards. These serpentine passages are disposed side-by-side, one adjacent the pressure surface and the other adjacent the suction surface on opposite sides of a main load carrying member which comprises a major part of the internal structure of the aerofoil.
Abstract translation: 用于燃气涡轮机翼型的内部流体冷却系统包括多个多通道冷却装置，每个多通道冷却装置由机翼内部的蛇形通道组成。 冷却流体 - 特别是空气 - 被供应到每个通道的入口端并通过多个排出孔排出以提供尖端，前缘，后缘和表面膜冷却。 第一蛇形通道的入口端定位成靠近前缘并向后流动，而第二蛇形通道的入口端靠近后缘定位并向前流动。 这些蛇形通道并排设置，一个邻近压力表面，另一个邻近主负荷承载构件的相对侧的吸入表面，该主负载承载构件包括机翼的内部结构的主要部分。
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