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公开(公告)号：US12085025B2
公开(公告)日：2024-09-10
申请号：US18387527
申请日：2023-11-07
申请人： RTX CORPORATION
CPC分类号： F02C7/36 , F02C3/04 , F02C3/06 , F02C3/107 , F02K3/025 , F02K3/06 , F05D2220/32 , F05D2220/323 , F05D2260/4031 , F05D2260/40311
摘要： A gas turbine engine includes, among other things, a propulsor section including a propulsor, a core engine, a gear arrangement that drives the propulsor. A compressor section includes a first compressor section and a second compressor section. A turbine section includes a first turbine and a second turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor section and a pressure ratio across the second compressor section, and greater than 40. The pressure ratio across the second compressor section is between 7 and 15, and the pressure ratio across the first compressor section is between 4 and 8.
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公开(公告)号：US12065965B1
公开(公告)日：2024-08-20
申请号：US18234340
申请日：2023-08-15
发明人： Victor L. Oechsle , Kenneth M. Pesyna , Bryan H. Lerg , Michael C. Monzella , Zachary A. Rauch , Michael Moser
摘要： A gas turbine engine includes a bypass duct and a rotating detonation augmentor. The bypass duct is configured to conduct air through a flow path arranged around an engine core of the gas turbine engine to provide thrust for propelling the gas turbine engine. The rotating detonation augmentor is located in the bypass duct and configured to be selectively operated to detonate fuel and a portion of the air to increase the thrust for propelling the gas turbine engine.
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公开(公告)号：US11988099B2
公开(公告)日：2024-05-21
申请号：US16900407
申请日：2020-06-12
IPC分类号： B64C11/00 , B64C11/18 , B64C11/46 , B64C11/48 , F01D9/02 , F01D17/16 , F02C6/20 , F02K1/46 , F02K3/02 , B64D27/00
CPC分类号： F01D17/16 , B64C11/001 , B64C11/18 , B64C11/46 , B64C11/48 , F01D9/02 , F02C6/206 , F02K1/46 , F02K3/025 , B64C11/00 , B64D2027/005 , F05B2240/12 , F05B2260/96 , F05D2220/324 , F05D2220/325 , F05D2250/30 , F05D2250/51 , F05D2250/52 , F05D2260/14 , Y02E10/72 , Y02T50/60
摘要： An unducted thrust producing system, includes a rotating element, a stationary element. An inlet may be located forward or aft of the rotating element and the stationary element. An exhaust may be located forward, aft, or between the rotating element and the stationary element.
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公开(公告)号：US11846238B2
公开(公告)日：2023-12-19
申请号：US17060171
申请日：2020-10-01
CPC分类号： F02C7/36 , F02C3/04 , F02C3/06 , F02C3/107 , F02K3/025 , F02K3/06 , F05D2220/32 , F05D2220/323 , F05D2260/4031 , F05D2260/40311
摘要： A turbofan gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10. A gear arrangement drives the fan section. A compressor section includes a low pressure compressor section and a high pressure compressor section. A turbine section drives the gear arrangement. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section, and greater than 40. The pressure ratio across the high pressure compressor section is between 7 and 15, and the pressure ratio across the low pressure compressor section is between 4 and 8.
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公开(公告)号：US10072584B2
公开(公告)日：2018-09-11
申请号：US14534318
申请日：2014-11-06
发明人： Michael R. Thomas , Nathan Snape
IPC分类号： F02K3/077 , F02C9/18 , F02K3/00 , F02C7/141 , F02C7/32 , F02K3/02 , F02C7/14 , F02K3/04 , F02C7/18
CPC分类号： F02C9/18 , F02C7/14 , F02C7/141 , F02C7/185 , F02C7/32 , F02K3/00 , F02K3/02 , F02K3/025 , F02K3/04 , F02K3/077
摘要： A gas turbine engine comprises a first bypass flow path housing configured within the engine, radially exterior to an engine core housing, and a second bypass flow path housing configured within the engine, radially exterior to the first bypass flow path housing. An axially downstream portion of the first bypass flow path housing includes a stepwise increase in area compared with an axially adjacent upstream portion of the first bypass flow path housing, thereby defining a component placement cavity in the axially downstream portion.
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公开(公告)号：US10001083B2
公开(公告)日：2018-06-19
申请号：US14335107
申请日：2014-07-18
申请人： MTU Aero Engines AG
发明人： Klaus Peter Rued , Werner Humhauser , Hermann Klingels , Rudolf Stanka , Eckart Heinrich , Hans-Peter Hackenberg , Stefan Weber , Claus Riegler , Erich Steinhardt , Jochen Gier , Manfred Feldmann , Norbert Huebner , Karl Maar
摘要： A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.
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公开(公告)号：US10001080B2
公开(公告)日：2018-06-19
申请号：US14462682
申请日：2014-08-19
申请人： The Boeing Company
发明人： David F. Cerra , Robert H. Willie
CPC分类号： F02K1/72 , F02K1/11 , F02K1/12 , F02K1/1207 , F02K1/70 , F02K1/763 , F02K3/025 , F05D2260/50
摘要： A thrust reverse variable area nozzle system for a nacelle of an aircraft engine system may include a reverse thrust opening disposed in the nacelle, and a thrust reverser door pivotally movable relative to the nacelle for selectively covering the reverse thrust opening, wherein the thrust reverser door is pivotally movable between a first position for completely covering the reverse thrust opening, a second position for partially uncovering a forward portion of the reverse thrust opening and discharging a bypass airflow through the forward portion of the reverse thrust opening in a forward direction, and a third position for partially uncovering an aft portion of the reverse thrust opening and discharging the bypass airflow through the aft portion of the reverse thrust opening in an aft direction.
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公开(公告)号：US09909506B2
公开(公告)日：2018-03-06
申请号：US14759979
申请日：2013-01-28
IPC分类号： F02C9/18 , F02K3/10 , F02K3/077 , F01D1/04 , F02K1/78 , F02K3/075 , F02C3/04 , F02C3/14 , F02C7/08 , F02K3/02
CPC分类号： F02C9/18 , F01D1/04 , F02C3/04 , F02C3/145 , F02C7/08 , F02K1/78 , F02K3/025 , F02K3/075 , F02K3/077 , F02K3/10 , F05D2220/32 , F05D2250/31
摘要： A gas turbine engine has a fan rotor for delivering air axially downstream into a core engine duct, which sequentially passes a turbine, a combustor, and a compressor. The core engine duct extends to a turning supply duct configured to turn the core air flow axially upstream so that the core air sequentially passes through the compressor, combustor and turbine, and into an exhaust conduit which turns the core airflow radially outwardly and axially downstream into an exhaust duct. A door is selectively opened to communicate a portion of the core airflow in the core engine duct to an augmentor with the exhaust duct isolated from the augmentor. The door is at a location prior to the core airflow reaching the compressor. A method is also disclosed.
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公开(公告)号：US09903274B2
公开(公告)日：2018-02-27
申请号：US14535614
申请日：2014-11-07
发明人： Carlos Enrique Diaz , William Dwight Gerstler , Michael Ralph Storage , Michael Thomas Kenworthy
IPC分类号： F02C1/00 , F02C7/14 , B64D33/12 , F02C7/18 , F28F27/02 , F28D1/047 , F28D7/08 , F02K3/02 , F28F13/00 , F02K3/06 , F28F1/12 , F28D21/00
CPC分类号： F02C7/14 , B64D33/12 , F02C7/18 , F02K3/025 , F02K3/06 , F05D2250/90 , F05D2300/505 , F28D1/047 , F28D7/08 , F28D2021/0021 , F28D2021/0026 , F28F1/12 , F28F13/00 , F28F27/02 , F28F2255/04 , F28F2280/10 , Y02T50/675
摘要： A heat exchanger apparatus including a surface cooler and a passive automatic retraction and extension system coupled to the surface cooler. The surface cooler having disposed therein one or more fluid flow channels configured for the passage therethrough of a heat transfer fluid to be cooled. The heat transfer fluid in a heat transfer relation on an interior side of said one or more fluid flow channels. The surface cooler including a plurality of fins projecting from an outer surface thereof. The passive automatic retraction and extension system including a thermal actuation component responsive to a change in temperature of at least one of the heat transfer fluid and a cooling fluid flow so as to actuate a change in a geometry of the surface cooler. Further disclosed is an engine including the heat exchanger apparatus.
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公开(公告)号：US20170122220A1
公开(公告)日：2017-05-04
申请号：US15411173
申请日：2017-01-20
IPC分类号： F02C7/36 , F02C7/06 , F02C9/18 , F01D25/24 , F01D5/06 , F01D9/02 , F01D21/00 , F02C3/06 , F02K3/06
CPC分类号： F02C7/36 , F01D5/06 , F01D5/12 , F01D9/02 , F01D21/003 , F01D25/24 , F01D25/28 , F02C3/04 , F02C3/06 , F02C3/107 , F02C7/06 , F02C7/28 , F02C9/18 , F02K3/025 , F02K3/06 , F05D2220/32 , F05D2220/323 , F05D2240/12 , F05D2240/35 , F05D2240/55 , F05D2260/40 , F05D2260/4031 , F05D2260/40311
摘要： A gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10 at cruise power, and a low fan pressure ratio of less than 1.45 measured across a fan blade alone. A gear arrangement drives the fan section. A compressor section includes both a low pressure compressor and a high pressure compressor. A guide vane includes a forward attachment, the forward attachment positioned aft of a plumbing connection area. A turbine section drives the gear arrangement, and may have a low pressure turbine with a low pressure turbine pressure ratio greater than 5:1, and a two stage high pressure turbine. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor and a pressure ratio across the high pressure compressor, and greater than 50, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the high pressure compressor is greater than 7.
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