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公开(公告)号：US10689987B2
公开(公告)日：2020-06-23
申请号：US16562701
申请日：2019-09-06
发明人： Thomas Veitch , Farid Abrari , Ernest Adique , Paul Aitchison , Daniel Fudge , Kari Heikurinen , Paul Stone , Tibor Urac
摘要： A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.
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公开(公告)号：US09303589B2
公开(公告)日：2016-04-05
申请号：US13687540
申请日：2012-11-28
发明人： Kari Heikurinen , Peter Townsend
CPC分类号： F04D29/384 , B23K20/129 , B23K2101/001 , F01D5/14 , F01D5/3061 , F01D5/34 , F02K3/06 , F04D29/023 , F04D29/321 , F04D29/324 , F04D29/329 , F04D29/34 , F04D29/388 , F04D29/644 , F05D2230/239 , F05D2260/40 , Y10T29/49245
摘要： A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (RHUB) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (RTIP) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (RHUB/RTIP) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.
摘要翻译： 一种用于涡轮风扇燃气涡轮发动机的风扇，风扇包括转子毂和多个径向延伸的与轮毂一体形成的叶片转子的径向延伸的风扇叶片。 每个风扇叶片定义一个前缘。 轮毂半径（RHUB）是轮毂相对于风扇中心线的前缘的半径。 尖端半径（RTIP）是风扇叶片尖端相对于风扇中心线的前缘的半径。 轮毂半径与顶端半径（RHUB / RTIP）的比率至少小于0.29。 在一个具体实施方案中，该比率在0.25和0.29之间。 在另一具体实施方案中，该比率小于或等于0.25。
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公开(公告)号：US10408223B2
公开(公告)日：2019-09-10
申请号：US15616302
申请日：2017-06-07
发明人： Kari Heikurinen , Peter Townsend
IPC分类号： F01D5/30 , B23K20/12 , F01D5/14 , F01D5/34 , F02K3/06 , F04D29/02 , F04D29/32 , F04D29/64 , F04D29/38 , F04D29/34 , B23K101/00
摘要： A fan for a turbofan gas turbine engine having a low hub-to-tip ratio is disclosed. The fan includes a rotor hub and a plurality of radially extending fan blades. Each fan blade defines a hub radius (RHUB), which is the radius of the leading edge at the hub relative to a centerline of the fan, and a tip radius (RTIP), which is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (RHUB/RTIP) is less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than 0.25.
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公开(公告)号：US09759230B2
公开(公告)日：2017-09-12
申请号：US14163588
申请日：2014-01-24
发明人： Kari Heikurinen , Ronald Dutton
CPC分类号： F04D29/541 , F01D5/143 , F01D5/145 , F04D19/028 , F04D29/547 , F04D29/681 , F05D2270/17
摘要： A multi-stage axial compressor with an inner wall including a step portion for each of the compressor stages. Each step portion is defined along a respective stage. Each step portion may extend over at least a majority of an axial length of the stage. Each step portion may optionally include a point aligned with a maximum thickness of the airfoil portions of the rotor blades and a point aligned with a maximum thickness of the stator vanes. Adjacent step portions are connected by a transition portion converging toward a central axis of the compressor from the upstream step to the downstream step. Each transition portion has a steeper slope than that of the adjacent step portions. A method of directing flow through a multi-stage axial flow compressor is also discussed.
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公开(公告)号：US10408231B2
公开(公告)日：2019-09-10
申请号：US15703472
申请日：2017-09-13
发明人： Thomas Veitch , Farid Abrari , Ernest Adique , Daniel Fudge , Kari Heikurinen , Paul Stone , Ignatius Theratil , Peter Townsend , Tibor Urac
摘要： A rotor for a gas turbine engine comprises a rotor having a hub and blades around the hub, and extending from the hub to tips. The tips include first and second tip portions between their respective tip leading edge and tip trailing edge. Tips are spaced from a rotational axis of the rotor by spans. A mean span of a first tip portion of a first blade is greater than a mean span of a corresponding first tip portion of a second blade. A mean span of a second tip portion the first blade is less than a mean span of a corresponding second tip portion of the second blade.
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公开(公告)号：US10370973B2
公开(公告)日：2019-08-06
申请号：US14725879
申请日：2015-05-29
发明人： Kari Heikurinen , Ron Dutton
摘要： A compressor airfoil of a gas turbine engine includes a pressure side and a suction side of the airfoil extending downstream from a stagnation point, the suction side including a suction side surface portion within a leading edge region, and a main suction side airfoil surface downstream from the suction side surface portion and extending contiguously therewith. The suction side surface portion having a compound curvature profile which includes at least a leading edge having a first curvature profile and a chamfered surface having a second curvature profile different from the first curvature profile. The chamfered surface being contiguous with and extending immediately downstream from the leading edge. The first curvature profile being curved. The second curvature profile of the chamfered surface being substantially flat and defining a substantially straight-line profile in a cross-section transverse to the span-wise axis of the airfoil.
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公开(公告)号：US11002293B2
公开(公告)日：2021-05-11
申请号：US15706311
申请日：2017-09-15
发明人： Karan Anand , Farid Abrari , Ernest Adique , Paul Aitchison , Daniel Fudge , Kari Heikurinen , Paul Stone , Tibor Urac , Thomas Veitch
摘要： A compressor rotor for a gas turbine engine includes a hub disposed about an axis of rotation and an outer surface forming a radially inner gaspath boundary, the outer surface defining a nominal hub diameter. A circumferential array of blades extends radially outwardly from the hub. A first inter-blade passage is defined between a first set of adjacent blades and has a first throat area. A second inter-blade passage is defined between a second set of adjacent blades and has a second throat area that is smaller than the first throat area. At least one scoop is disposed in the second inter-blade passage, the scoop defining a cavity extending radially into the outer surface of the hub relative to the nominal hub diameter.
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公开(公告)号：US10443411B2
公开(公告)日：2019-10-15
申请号：US15707133
申请日：2017-09-18
发明人： Thomas Veitch , Farid Abrari , Ernest Adique , Paul Aitchison , Daniel Fudge , Kari Heikurinen , Paul Stone , Tibor Urac
摘要： A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.
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