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公开(公告)号：US11738877B2
公开(公告)日：2023-08-29
申请号：US16832607
申请日：2020-03-27
Applicant: ROLLS-ROYCE plc
Inventor： Steven A Radomski , Richard G Stretton
CPC classification number: B64D27/26 , B64C7/00 , F01D11/005 , F16J15/102 , B64D2027/262 , F05D2240/55
Abstract: A seal assembly for a gas turbine engine having a rotor arranged to rotate about an axis in use. The seal assembly has a static support structure for the gas turbine engine and a casing structure of the engine. Rotation of the engine rotor causes a deflection of the casing structure relative to the static support structure in a first direction. A seal is provided at an interface between the static support structure and the casing structure, and comprising a first seal portion and a second seal portion spaced from one another in the first direction. The first seal portion is provided against a first surface of the casing structure and the second seal portion is provided against a second surface of the casing structure opposing the first surface. In an at-rest state in which the engine is not operational, the first and second surfaces are offset from an equilibrium position with respect to the static support structure such that there is a difference in compression of the first seal portion and the second seal portion between the static support structure and the casing structure. The offset is in a direction opposite to the first direction.
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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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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：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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Patent Agency Ranking
公开(公告)号：US11945595B2
公开(公告)日：2024-04-02
申请号：US18074270
申请日：2022-12-02
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton
CPC classification number: B64D27/40 , B64D27/10 , B64D27/402 , F02C7/20 , B64D27/406
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：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：US11187109B2
公开(公告)日：2021-11-30
申请号：US16909060
申请日：2020-06-23
Applicant: ROLLS-ROYCE plc
Inventor： Richard G Stretton , Steven A Radomski
IPC: F01D25/24
Abstract: A gas turbine engine casing is described as having a cowl door hinged to a casing support structure by at least one hinge. The cowl door is openable outwardly from the casing to expose a casing interior. The hinge is located above a longitudinal axis of the casing and comprises a pivoting linkage arranged such that, upon actuation between closed and open cowl door conditions, the pivoting linkage moves an upper portion of the cowl door downwards towards the longitudinal axis.
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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：US11421592B2
公开(公告)日：2022-08-23
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
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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