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公开(公告)号：US20220056916A1
公开(公告)日：2022-02-24
申请号：US17520126
申请日：2021-11-05
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Michael C. WILLMOT
Abstract: A gas turbine engine for an aircraft includes an engine core having a core length and comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius, wherein a ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in a range from 1.2 to 2.0; and wherein the engine core length is in a range from 150 cm to 320 cm.
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公开(公告)号：US20200049022A1
公开(公告)日：2020-02-13
申请号：US16507231
申请日：2019-07-10
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Steven A. RADOMSKI
Abstract: A gas turbine engine (10) comprises a bypass duct cowl (21), an engine core housing (22) defining an engine core inlet, a bypass fan (13) and a plurality of outlet guide vanes (24). Each outlet guide vane 24 extends between a radially inner surface of the bypass duct cowl (21) and a radially outer surface of the engine core housing (22, 23) to define an outlet guide vane span (SPANOGV). The outlet guide vanes (24) are configured to support the engine core housing (22, 23) relative to the bypass duct cowl (21). The bypass fan (13) and an engine core inlet (34) define a bypass ratio between 10 and 17, and a ratio of the outlet guide vane span (OGVSPAN) to a bypass fan radius (RFAN) is between 0.45 and 0.55.
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公开(公告)号：US20200011250A1
公开(公告)日：2020-01-09
申请号：US16437907
申请日：2019-06-11
Applicant: ROLLS-ROYCE plc
Inventor： Alan R. MAGUIRE , Richard G. STRETTON
Abstract: Apparatus for a gas turbine engine, the apparatus comprising: a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing; a plurality of guide vanes extending radially from the outer wall of the core engine casing; a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; and an accessory gear box positioned within the second cavity of the torque box.
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公开(公告)号：US20210239051A1
公开(公告)日：2021-08-05
申请号：US17235390
申请日：2021-04-20
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Tim O'HANRAHAN
Abstract: A gas turbine engine comprising a planetary gear train, and a core engine casing. The gear train has a ratio of greater than approximately 3.0, with an input to the gear train being operatively connected to the compressor section, and an output from the gear train being operatively connected to the fan. The core engine casing encloses the compressor section and the turbine section. The fan has a diameter F, and the core engine casing has a diameter C. The core engine casing diameter C varies along an axial length of the core engine casing, and a ratio (C/F) of the core engine casing diameter C to the fan diameter F is within the range 0.2
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公开(公告)号：US20200347848A1
公开(公告)日：2020-11-05
申请号：US16825504
申请日：2020-03-20
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Michael C. WILLMOT
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.
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Patent Agency Ranking
公开(公告)号：US20200248699A1
公开(公告)日：2020-08-06
申请号：US16825361
申请日：2020-03-20
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Michael C. WILLMOT
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.
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公开(公告)号：US20170292473A1
公开(公告)日：2017-10-12
申请号：US15459724
申请日：2017-03-15
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON
Abstract: An aircraft gas turbine engine nacelle comprises a thrust reversal arrangement. The thrust reversal arrangement comprises at least first and second circumferentially spaced fixed thrust reverser cascade boxes each comprising a plurality of thrust reverser vanes configured to direct air forwardly and circumferentially and at least one inter-leaved translating circumferential turning vane configured to direct air in a direction having a circumferential component. The circumferential turning vane is moveable from a stowed position provided between the first and second circumferentially spaced thrust reverser cascade boxes, and a deployed position axially rearwardly of the thrust reverser cascade boxes.
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公开(公告)号：US20220135234A1
公开(公告)日：2022-05-05
申请号：US17502203
申请日：2021-10-15
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON
Abstract: A gas turbine engine for mounting to an airframe of an aircraft comprises an engine core; a fan located upstream of the engine core; a bifurcation spanning a bypass duct defined between the engine core and a nacelle surrounding the gas turbine engine, the bifurcation comprising aerodynamically shaped fairings defining an interior space therebetween; and a cabin blower system arranged in the interior space of the upper bifurcation.
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公开(公告)号：US20190309688A1
公开(公告)日：2019-10-10
申请号：US16351841
申请日：2019-03-13
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Tim O'HANRAHAN
Abstract: A gas turbine engine comprising a planetary gear train, and a core engine casing. The gear train has a ratio of greater than approximately 3.0, with an input to the gear train being operatively connected to the compressor section, and an output from the gear train being operatively connected to the fan. The core engine casing encloses the compressor section and the turbine section. The fan has a diameter F, and the core engine casing has a diameter C. The core engine casing diameter C varies along an axial length of the core engine casing, and a ratio (C/F) of the core engine casing diameter C to the fan diameter F is within the range 0.2
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公开(公告)号：US20190300190A1
公开(公告)日：2019-10-03
申请号：US16353210
申请日：2019-03-14
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON
Abstract: A gas turbine engine (10) for an aircraft (90) comprises an engine core (11, 60) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; an engine casing (61) arranged to at least partially surround the engine core (11), the engine casing comprising at least one first elongate formation (67) on its outer surface, the first elongate formation extending in an axial direction; and a housing (55) arranged to surround the engine core (11, 60), the housing (55) comprising at least one second elongate formation (57) on an inner surface of the housing, the second elongate formation (57) extending in the axial direction. The engine casing (61) is detachably connected to the housing (55) by interengagement of the first and second elongate formations (57, 67).
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