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公开(公告)号：US20250012234A1
公开(公告)日：2025-01-09
申请号：US18746600
申请日：2024-06-18
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON
Abstract: A propulsion device comprises: a casing structure to surround a fan of the propulsion device; a core structure to support a core of the propulsion device; a thrust reverser unit comprising two thrust reverser halves, each thrust reverser half being pivotable about a respective hinge line between an open position for access to the core structure, and a closed position. Upper and lower support members extend from the casing structure at diametrically opposing sides of a centreline axis. For each thrust reverser half there is at least one locating arrangement comprising an upper locating arrangement with cooperating portions configured to engage each other as the thrust reverser half moves towards the closed position in a closing operation; and/or a lower locating arrangement with cooperation portions configured to engage each other as the thrust reverser half moves towards the closed position in the closing operation.
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公开(公告)号：US20230323810A1
公开(公告)日：2023-10-12
申请号：US18184008
申请日：2023-03-15
Applicant: ROLLS-ROYCE plc
Inventor： Christopher A. MURRAY , Nicholas HOWARTH , Richard G. STRETTON
CPC classification number: F02C6/08 , B64D13/06 , B64D2013/0618 , F05D2260/205 , F05D2220/323
Abstract: There is provided a gas turbine engine comprising a blower system for supplying pressurised air to an airframe via an airframe port. The blower system comprises a compressor configured to receive air from a bypass duct or a core of the gas turbine engine and to discharge compressed air into a delivery line extending from the compressor to the airframe port. The blower system also comprises a heat exchanger configured to transfer heat from the compressed air to a coolant and a valve arrangement configured to switch between operation of the blower system in a baseline mode and a cooling mode, the valve arrangement being configured to: selectively divert the compressed air within the delivery line to the heat exchanger for operation in the cooling mode; and/or selectively provide the coolant to the heat exchanger for operation in the cooling mode.
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公开(公告)号：US20190382122A1
公开(公告)日：2019-12-19
申请号：US16416708
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 comprises a pylon attachment, a shaft defining an engine centreline which lies in an engine central plane intersecting the pylon attachment, a fan defining a fan plane normal to the engine centreline and an intake upstream of the fan plane. The geometric centreline of the intake coincides with the engine centreline at an axial position corresponding to the downstream end of the intake and curves away from the engine centreline upstream of said axial position. The engine may be mounted on one side of an aircraft such that the orientation of the highlight plane of the intake is aligned to the air flow field of the aircraft on that side during flight.
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公开(公告)号：US20220065171A1
公开(公告)日：2022-03-03
申请号：US17401647
申请日：2021-08-13
Applicant: ROLLS-ROYCE plc
Inventor： Richard G. STRETTON , Richard SHARPE
Abstract: A gas turbine engine includes an engine core including a compressor, a combustor, and a turbine, the compressor being connected to the turbines through a respective shaft; and a cabin blower system comprising: an electric variator comprising a first electrical machine connected to a first shaft arranged along a first axis, a second electrical machine connected to a second shaft arranged along a second axis, and a power management system; a cabin blower comprising a compressor driven by a third shaft arranged along a third axis, the compressor comprising an air inlet and an air outlet; and a differential gearbox. The gas turbine engine further includes an accessory gearbox arranged within an accessory gearbox casing and adapted to drive the cabin blower system.
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公开(公告)号：US20210301827A1
公开(公告)日：2021-09-30
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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Patent Agency Ranking
公开(公告)号：US20210164478A1
公开(公告)日：2021-06-03
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：US20210148306A1
公开(公告)日：2021-05-20
申请号：US17146910
申请日：2021-01-12
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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公开(公告)号：US20210079810A1
公开(公告)日：2021-03-18
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：US20200347803A1
公开(公告)日：2020-11-05
申请号：US16929806
申请日：2020-07-15
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：US20200317354A1
公开(公告)日：2020-10-08
申请号：US16832607
申请日：2020-03-27
Applicant: ROLLS-ROYCE plc
Inventor： Steven A. RADOMSKI , Richard G. STRETTON
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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