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公开(公告)号：US20240360802A1
公开(公告)日：2024-10-31
申请号：US18644812
申请日：2024-04-24
CPC分类号： F02K1/766 , F02K1/72 , F05D2220/323
摘要： An aircraft nacelle equipped with a thrust reversal device which comprises: at least one deflection system configured to deflect an air stream channeled in the nacelle toward a lateral opening of the nacelle, in the activated state, at least one orientation system having: at least one transverse deflector configured to orient the air stream deflected by the deflection system in radial directions and toward the upstream end of the nacelle, at least one longitudinal deflector configured to orient the air stream deflected by the deflection system in at least one direction that forms an angle with a radial direction.
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公开(公告)号：US20240360801A1
公开(公告)日：2024-10-31
申请号：US18644977
申请日：2024-04-24
CPC分类号： B64D29/06 , F02K1/32 , F05D2220/323
摘要： An aircraft nacelle equipped with a thrust reversal device which comprises: at least one deflection system configured to deflect an air stream channeled in the nacelle toward a lateral opening of the nacelle, in the activated state; at least one orientation system having: at least one transverse deflector configured to orient the air stream deflected by the deflection system toward the upstream end of the nacelle, at least one downstream deflector positioned at the downstream edge of the lateral opening and configured to deflect the air stream deflected by the deflection system toward the upstream end of the nacelle.
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公开(公告)号：US20240360767A1
公开(公告)日：2024-10-31
申请号：US18687378
申请日：2022-08-31
发明人： Jean-Hilaire LEXILUS
IPC分类号： F01D5/22
CPC分类号： F01D5/22 , F05D2220/323 , F05D2230/60 , F05D2240/30 , F05D2260/96
摘要： A damper for an impeller of an aircraft turbomachine includes an elastic member housed in two metallic half-shells so as to allow a relative movement of the half-shells relative to one another.
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公开(公告)号：US20240359813A1
公开(公告)日：2024-10-31
申请号：US18770884
申请日：2024-07-12
申请人： RTX CORPORATION
摘要： An aircraft propulsion system includes at least two gas turbine engines and a bottoming cycle system where a working fluid is circulated within a closed circuit that includes a bottoming compressor section and a bottoming turbine section. Each of the at least two gas turbine engines include a primary heat exchanger for communicating thermal energy into the working fluid of the bottoming cycle.
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公开(公告)号：US20240352891A1
公开(公告)日：2024-10-24
申请号：US18749076
申请日：2024-06-20
申请人： ROLLS-ROYCE plc
CPC分类号： F02C7/06 , F01D25/16 , F02K3/04 , F02C7/36 , F05D2200/14 , F05D2220/323 , F05D2240/30 , F05D2240/50
摘要： A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.
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公开(公告)号：US12123356B2
公开(公告)日：2024-10-22
申请号：US17379233
申请日：2021-07-19
发明人： Stephen G. Pixton , Karl L. Hasel
CPC分类号： F02C7/36 , F05D2220/323 , F05D2260/40311
摘要： A gas turbine engine includes a fan drive turbine driving a low pressure compressor, and driving a gear reduction to in turn drive a fan rotor at a speed slower than the fan drive turbine. The turbine section further includes a high pressure turbine driving the high pressure compressor. The fan drive turbine and low pressure compressor are connected by a shaft and the fan drive turbine, the shaft and the low pressure compressor define a low pressure spool. The gas turbine engine is rated to provide an amount of thrust at maximum takeoff, and a low spool thrust ratio defined as a ratio of a torque on the low pressure spool at maximum takeoff in ft-lbs and the maximum takeoff thrust being defined in lbf, with the low spool torque ratio being greater than or equal to 0.70 ft-lb/lbf, and less than or equal to 1.2 ft-lb/lbf.
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公开(公告)号：US12123313B2
公开(公告)日：2024-10-22
申请号：US17366654
申请日：2021-07-02
申请人： RTX Corporation
发明人： Kevin Todd Lowe , Gwibo Byun , Albert P. Krejmas
CPC分类号： F01D21/003 , B64D27/18 , B64D43/00 , F02C9/00 , G01F1/661 , F05D2220/323 , F05D2260/80 , F05D2270/3061 , F05D2270/804
摘要： An aircraft includes a gas turbine engine and an optically-based measurement system. The gas turbine engine is configured to ingest a first mass flow and to exhaust a second mass flow. The optically-based measurement system is configured to determine the first and second mass flows in response to performing an imaging process on the gas turbine engine.
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公开(公告)号：US12123310B2
公开(公告)日：2024-10-22
申请号：US18480627
申请日：2023-10-04
申请人： RTX CORPORATION
发明人： Stephen G. Pixton , Gary Collopy , Ozhan Turgut
CPC分类号： F01D15/12 , F02C3/00 , F02C7/32 , F05D2220/323
摘要： A gas turbine engine according to an example of the present disclosure may include, among other things, a propulsor section including a propulsor having a plurality of propulsor blades, a geared architecture, a first spool including a first shaft that interconnects a first compressor and a first turbine, the first driving the propulsor through the geared architecture. The gas turbine engine is rated to provide an amount of thrust at ground idle, and the gas turbine engine is rated to provide an amount of thrust at maximum takeoff. A thrust ratio is defined as a ratio of the amount of thrust at ground idle divided by the amount of thrust at maximum takeoff. The thrust ratio can be less than or equal to 0.050.
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公开(公告)号：US12110844B2
公开(公告)日：2024-10-08
申请号：US17987671
申请日：2022-11-15
申请人： ROHR, INC.
IPC分类号： F02K1/82
CPC分类号： F02K1/827 , F05D2220/323 , F05D2260/96
摘要： A center plug for attenuating noise in a gas turbine engine includes an inner skin, a forward bulkhead, and aft bulkhead, and a noise attenuation panel. The inner skin has a substantially cylindrical shape and extending along an axial centerline. The forward bulkhead is disposed proximate a forward end of the inner skin. The forward bulkhead is connected to and extends radially outward from the inner skin. The aft bulkhead is disposed proximate an aft end of the inner skin. The aft bulkhead is connected to and extending radially outward from the inner skin. The noise attenuation panel is positioned intermediate the inner skin and partially divides a region bounded by the inner skin, the forward bulkhead and the aft bulkhead into.
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公开(公告)号：US12110827B1
公开(公告)日：2024-10-08
申请号：US18311660
申请日：2023-05-03
CPC分类号： F02C7/236 , F02C7/32 , F05D2220/323 , F05D2270/07
摘要： A turbo engine for an aircraft includes a gas turbine engine having a combustion section and a fuel system to provide pressurized fuel to the combustion section. The fuel system includes a fuel tank and a boost pump and a main pump driven by an accessory gearbox that is powered by a shaft of the turbo engine. The boost pump and the main pump are arranged in series. The fuel system also includes an auxiliary pump and a controller to operate the fuel system in (1) a first mode in which the boost pump and the main pump produce pressurized fuel for the turbo engine and the auxiliary pump is deactivated, and (2) a second mode in which the auxiliary pump is activated and produces the pressurized fuel for the turbo engine while bypassing the boost pump and the main pump.
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