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公开(公告)号：US11629666B2
公开(公告)日：2023-04-18
申请号：US17184903
申请日：2021-02-25
发明人： Aviad Brandstein , Avi Ponchek
摘要： A method for converting a turbofan engine including providing a turbofan engine and converting the turbofan engine. The turbofan engine includes a core engine (including at least one high pressure spool assembly and a combustion chamber), and an unmodified fan configured for providing at least a bypass flow bypassing the core engine, the fan being mechanically coupled to a low pressure turbine that is in turn driven by the core engine. The conversion includes modifying or replacing the unmodified fan to provide a modified fan, the modified fan configured for generating a reduced bypass flow with respect to said fan bypass flow during operation of the converted turbofan engine corresponding to at least one set of engine conditions, enabling said low pressure turbine to generate an excess shaft power above a baseline shaft power required for driving the modified fan during operation of the converted turbofan engine.
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公开(公告)号：US11623723B2
公开(公告)日：2023-04-11
申请号：US17022336
申请日：2020-09-16
发明人： John Cromie , Carey Hijmans
摘要： Aspects of the technology relate to a propeller blade assembly that is used in lateral propulsion systems for lighter-than-air high altitude platforms designed to operate, e.g., in the stratosphere. During operation, the propeller of the assembly is pointed along a specified heading and rotates at a selected velocity (e.g., hundreds or thousands of revolutions per minute). Power is supplied to the propeller as needed during lateral propulsion to move the platform along a particular trajectory or to remain on station over a given geographic location. In certain circumstances, the propeller may become damaged. This can include one or more blades breaking or shattering, which can result in failure of the propeller and potentially the entire LTA platform. The technology provides blades that are sufficiently flexible to avoid breakage or shattering due to debris impact or envelope entanglement, or otherwise shed a load. This can avoid catastrophic failure during stratospheric operation.
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公开(公告)号：US11619134B2
公开(公告)日：2023-04-04
申请号：US16326626
申请日：2016-08-22
发明人： Koji Wakashima , Makoto Ozaki , Kimio Hatano , Bumpei Hashimoto , Hiroshi Saito , Takuya Rikimaru
摘要： A mixed-flow turbine wheel includes: a plurality of rotor blades disposed on a circumferential surface of the hub at intervals in a circumferential direction and configured such that each of the plurality of rotor blades has a leading edge which includes, in a meridional view, an oblique edge portion where a distance between the leading edge and an axis of the rotational shaft decreases from a tip side toward a hub side, and a sensor detection surface having a flat shape and being applied with a marking which is detectable by an optical sensor device. The sensor detection surface is formed on at least one of the circumferential surface of the hub or an edge portion of a reference rotor blade being one of the plurality of rotor blades, such that, in the meridional view, a trailing-edge side angle of two angles formed between the axis of the rotational shaft and a normal of the sensor detection surface is smaller than a trailing-edge side angle of two angles formed between the axis of the rotational shaft and a normal of the oblique edge portion.
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公开(公告)号：US11603762B2
公开(公告)日：2023-03-14
申请号：US16438134
申请日：2019-06-11
发明人： Martin Babak
摘要： An exhaust turbocharger turbine wheel can include a hub that includes a nose, a backdisk with a shaft joint portion, and a rotational axis; blades that extend from the hub to define exhaust flow channels where each of the blades includes a leading edge, a trailing edge, a hub profile, a shroud profile, a pressure side, and a suction side; where the backdisk includes an outer perimeter radius measured from the rotational axis of the hub, an intermediate radius at the shaft joint portion measured from the rotational axis of the hub, and an annular recess disposed between the intermediate radius and the outer perimeter radius and defined in part by three-dimensional bolster regions, where each of the three-dimensional bolster regions includes a footprint and a height measured at least in part in a direction of the rotational axis of the hub.
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公开(公告)号：US11598212B2
公开(公告)日：2023-03-07
申请号：US16758576
申请日：2018-10-25
摘要： A method for balancing a set of blades intended to be arranged on a bare disc of an aircraft engine, the bare disc comprising a defined number of numbered cells (ai) intended to receive the same defined number of blades, which can have a spread of mass, the method comprising the following steps:—sorting the blades by monotonic order of their mass (mi) to form an ordered set of blades,—separating the ordered set of blades in a balanced manner into four lobes constituted by a first large lobe, by a second large lobe, by a first small lobe and by a second small lobe, the blades being classified into each lobe according to a current placement order, and—arranging the four lobes on the bare disc by making the current placement order of the blades correspond to the numbered cells of the bare disc.
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公开(公告)号：US11577820B2
公开(公告)日：2023-02-14
申请号：US17203862
申请日：2021-03-17
申请人： Ratier-Figeac SAS
发明人： Bruno Seminel
摘要： A propeller blade arrangement comprising a propeller blade attached to and rotatable with a hub, via a retention bearing, the blade being rotatable about a center line of the blade, the retention bearing configured to tilt the blade such that its center line is tilted with respect to the hub.
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公开(公告)号：US20230042974A1
公开(公告)日：2023-02-09
申请号：US17969179
申请日：2022-10-19
摘要： A gas turbine engine includes, among other things, a propulsor section including a propulsor hub, the hub including a hub diameter supporting a plurality of propulsor blades. A compressor section includes a first compressor and a second compressor. A turbine section includes a first turbine and a second turbine. A geared architecture interconnects the first turbine and the propulsor hub. The geared architecture includes a gear volume. A compressor inlet passage is disposed annularly about the geared architecture.
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公开(公告)号：US11549373B2
公开(公告)日：2023-01-10
申请号：US17123344
申请日：2020-12-16
发明人： Michael G. McCaffrey
摘要： A turbine section for a gas turbine engine according to an example of the present disclosure includes, among other things, a first turbine rotor coupled to a first turbine shaft. The first turbine shaft is rotatable about a longitudinal axis. A second turbine rotor is coupled to a second turbine shaft. The second turbine shaft is rotatable about the longitudinal axis, and the second turbine rotor is axially aft of the first turbine rotor relative to the longitudinal axis. An aft bearing assembly rotatably supports the second turbine shaft. The second turbine rotor includes a disk assembly that carries at least one row of turbine blades. The disk assembly is mechanically attached to the second turbine shaft at an attachment point. The attachment point is axially aft of the aft bearing assembly such that an aft portion of the second turbine shaft is cantilevered from the aft bearing system with respect to the longitudinal axis. The disk assembly includes a metallic material. Each of the turbine blades comprises a ceramic matrix composite (CMC) material.
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公开(公告)号：US11541998B2
公开(公告)日：2023-01-03
申请号：US16787845
申请日：2020-02-11
申请人： MagLev Aero Inc.
发明人： Ian Morris Randall
IPC分类号： B64C27/20 , H02K1/2786 , H02N15/00 , F01D5/02 , B64C27/10 , B64D27/24 , H02K1/17 , H02K1/2793 , B64C29/00 , B64C13/50 , H02K21/22 , B64C27/32 , H02K1/18 , H02K11/21 , B64C27/473 , H02K21/24 , B64C13/26 , B64C27/72 , B64D35/02 , B64D35/04 , H02K1/28 , H02K16/00 , B64C27/68 , B64C27/00 , B64C29/02 , B64C27/14 , B64C27/80 , B64C39/02 , B64D35/06
摘要： Systems and methods relate to a vertical takeoff and landing (VTOL) platform that can include a stator and a rotor magnetically levitated by the stator. The rotor and stator can be annular, such that the rotor rotates about a rotational axis. The stator can include magnets that provide guidance, levitation, and drive forces to drive the rotor, as well as to control operation of rotor blades of the rotor that can be independently rotated to specific pitch angles to control at least one of lift, pitch, roll, or yaw of the VTOL platform. Various controllers can be used to enable independent and redundant control of components of the VTOL platform.
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10.发明申请公开(公告)号：US20220403741A1
公开(公告)日：2022-12-22
申请号：US17891312
申请日：2022-08-19
发明人： Nicholas Joseph Kray , Nitesh Jain
IPC分类号： F01D5/02
摘要： A rotor blade for a gas turbine engine includes an airfoil section and a root section extending along a longitudinal direction between an upstream surface and a downstream surface. The root section further extends along a radial direction between an inner surface positioned at an inner end of the root section and an outer end coupled to the airfoil section. Moreover, the root section extends along a circumferential direction between a first side surface and a second side surface. Furthermore, the root section defines a longitudinal centerline extending along the longitudinal direction and positioned equidistant from the inner surface and the outer end in the radial direction. The root section includes a first portion formed from a composite material and a second portion formed from a metallic material, with the longitudinal centerline extending through the second portion of the root section.
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