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公开(公告)号：US11988169B2
公开(公告)日：2024-05-21
申请号：US17175092
申请日：2021-02-12
Applicant: ROLLS-ROYCE PLC
Inventor： Richard G Stretton , Michael C Willmot
CPC classification number: F02K3/06 , F02C3/04 , F02C7/04 , F02C7/36 , B64D27/12 , B64D2033/0286 , F05D2220/32 , F05D2240/24 , F05D2260/80 , Y02T50/60
Abstract: A gas turbine engine for an aircraft having an engine core configured with a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, the fan comprising a plurality of fan blades, with a nacelle surrounding the gas turbine engine, and a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle. An outer wall axis is defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle. An outer bypass duct wall angle is defined as the angle between the outer wall axis and the centreline, and the outer bypass duct wall angle is in a range from −15 to −2.5 degrees.
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公开(公告)号：US11053947B2
公开(公告)日：2021-07-06
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：US12163464B2
公开(公告)日：2024-12-10
申请号：US17749908
申请日：2022-05-20
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Michael C Willmot , Nicholas Grech
Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass exhaust nozzle pressure ratio core exhaust nozzle pressure ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.-
公开(公告)号：US10981663B2
公开(公告)日：2021-04-20
申请号：US16934767
申请日：2020-07-21
Applicant: ROLLS-ROYCE PLC
Inventor： Richard G Stretton , Michael C Willmot
Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: the turbine diameter at an axial location of the lowest pressure rotor stage a distance f rom a ground plane to the wing and wherein an engine blockage ratio of: ( 2 × the fan tip radius/the engine length ) the downstream blockage ratio is in the range from 2.5 to 4.
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公开(公告)号：US11408428B2
公开(公告)日：2022-08-09
申请号：US17174967
申请日：2021-02-12
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：US11339713B2
公开(公告)日：2022-05-24
申请号：US16398742
申请日：2019-04-30
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Michael C Willmot , Nicholas Grech
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass exhaust nozzle pressure ratio core exhaust nozzle pressure ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.-
公开(公告)号：US11204037B2
公开(公告)日：2021-12-21
申请号：US17338159
申请日：2021-06-03
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Michael C Willmot
Abstract: A gas turbine engine for an aircraft includes an engine core including an engine core, 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. 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. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.
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公开(公告)号：US10760530B2
公开(公告)日：2020-09-01
申请号：US16423541
申请日：2019-05-28
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Michael C Willmot
Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the maximum take - off rotational speed of the fan fan - turbine radius difference ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.
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公开(公告)号：US10648475B1
公开(公告)日：2020-05-12
申请号：US16671215
申请日：2019-11-01
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：US10583932B1
公开(公告)日：2020-03-10
申请号：US16398909
申请日：2019-04-30
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Michael C Willmot
Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is: the turbine diameter at an axial location of the lowest pressure rotor stage ground plane to wing distance and a quasi-non-dimensional mass flow rate Q defined as: Q = W T 0 P 0 · A flow where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; and Aflow is the flow area of the fan face in m2, and wherein a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 to 0.01.
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