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公开(公告)号：US20230182911A1
公开(公告)日：2023-06-15
申请号：US18061330
申请日：2022-12-02
Applicant: ROLLS-ROYCE plc
Inventor： Richard G STRETTON
CPC classification number: B64D27/26 , B64D27/10 , B64D33/08 , F02C7/20 , B64D2027/268 , F05D2220/323 , F05D2240/90
Abstract: A gas turbine engine includes a support structure for attaching the engine to an aircraft pylon. The support structure includes: an engine-side interface member, a pylon-side interface member interfacing to the engine-side interface member, and a top V-shaped connection formation above the engine core and pair of side V-shaped connection formations on opposite lateral sides of the engine core, each V-shaped connection formation being formed by a pair of connection members meeting at a vertex, the vertex of the top V-shaped connection formation joining to the top of the engine-side interface member, the vertices of the side V-shaped connection formations respectively joining to the bottom ends of the engine-side interface member, and the connection members extending forwardly from their respective vertices to join to front fixation points at the core casing.
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公开(公告)号：US20230182912A1
公开(公告)日：2023-06-15
申请号：US18074270
申请日：2022-12-02
Applicant: ROLLS-ROYCE plc
Inventor： Richard G STRETTON
CPC classification number: B64D27/26 , B64D27/10 , F02C7/20 , B64D2027/268
Abstract: A support structure for attaching an engine to an aircraft pylon at front, mid and rear attachment positions thereof, including a front mount joined to the engine and configured to attach to the pylon at the front attachment position and a rear mount joined to a core casing to attach to the pylon at the rear attachment position, each of the front and rear mounts configured to transfer lateral and vertical loads from the engine to the pylon, and the rear mount being spaced from the front mount such that yaw and pitch torques are transferred from the engine to the pylon through the front and rear mounts. The support structure also includes an axial load transfer formation to transfer axial loads from the engine to the pylon and a roll-torque transfer formation to transfer roll torque from the core casing to the pylon.
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公开(公告)号：US20210164417A1
公开(公告)日：2021-06-03
申请号：US17175092
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 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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Patent Agency Ranking
公开(公告)号：US20200023984A1
公开(公告)日：2020-01-23
申请号：US16460307
申请日：2019-07-02
Applicant: ROLLS-ROYCE PLC
Inventor： Chia Hui LIM , Richard G STRETTON , Christopher T J SHEAF
Abstract: A mounting arrangement for mounting an aircraft gas turbine engine to an aircraft includes an engine nacelle with a distal assembly including a part annular engine cowl, a gas turbine engine core housing surrounded by the cowl and a distal bifurcation extending between the engine core housing and engine cowl in a first direction to define a first axis. The mounting arrangement includes a proximal assembly having a mount configured to mount the proximal assembly to the engine core housing. The proximal assembly includes a pylon configured to mount the proximal assembly to mounting location such as a wing of the aircraft at an engine mounting location. The pylon extends in a line between the wing and the engine core housing to define a second axis which is normal to a distal surface of the wing at the engine mounting location and is non-parallel to the vertical axis.
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公开(公告)号：US20240376847A1
公开(公告)日：2024-11-14
申请号：US18781563
申请日：2024-07-23
Applicant: ROLLS-ROYCE plc
Inventor： Richard G STRETTON , Michael C WILLMOT
Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.
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公开(公告)号：US20230028367A1
公开(公告)日：2023-01-26
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：US20200063604A1
公开(公告)日：2020-02-27
申请号：US16544996
申请日：2019-08-20
Applicant: ROLLS-ROYCE plc
Inventor： Richard G STRETTON , David WESTON , Nicholas P ROSE
IPC: F01D25/24
Abstract: There is disclosed a gas turbine engine for an aircraft comprising: a propulsive fan having a plurality of fan blades; a fan casing; and an air intake; wherein the air intake is mechanically coupled to the fan casing at a point having an axial position that is within a range of axial positions from a first axial position that is rearward of the leading edge of the fan blade at its radial tip by an axial component of a blade chord length, to a second axial position that is forward of the leading edge of the fan blade tip by an axial component of the blade chord length.
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公开(公告)号：US20190383215A1
公开(公告)日：2019-12-19
申请号：US16416766
申请日：2019-05-20
Applicant: ROLLS-ROYCE plc
Inventor： Christopher T J SHEAF , Richard G STRETTON , Chia Hui LIM
IPC: F02C7/04
Abstract: A gas turbine engine (100) for an aircraft comprises a pylon attachment (112) and a shaft (108) defining an engine centreline (110). The engine centreline lies in an engine central plane (120) which intersects the pylon attachment. The gas turbine engine comprises an intake (104) having a non-axisymmetric geometry and a medial plane (130) defining left and right halves of the intake. The left and right halves are configured for at least one of optimum cross wind performance, optimum incidence performance and optimum cruise performance when the medial plane is aligned with a vertical plane. The intake is installed so that the medial plane is angularly offset with respect to the engine central plane. The engine may be installed on a wing of an aircraft with the medial plane closer to its optimal orientation than is the case for a conventional engine.
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