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公开(公告)号：US12276225B2
公开(公告)日：2025-04-15
申请号：US18184008
申请日：2023-03-15
Applicant: ROLLS-ROYCE plc
Inventor： Christopher A. Murray , Nicholas Howarth , Richard G. Stretton
IPC: F02C6/08 , B64D13/00 , B64D13/02 , B64D13/08 , B64D33/08 , F02C7/12 , F02C7/18 , F02C9/16 , F02C9/20 , F02C9/54
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：US20250035037A1
公开(公告)日：2025-01-30
申请号：US18913201
申请日：2024-10-11
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 2.0 under aircraft cruise conditions.-
公开(公告)号：US10968835B2
公开(公告)日：2021-04-06
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：US11525407B2
公开(公告)日：2022-12-13
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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Patent Agency Ranking
公开(公告)号：US10539095B2
公开(公告)日：2020-01-21
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：US12258908B2
公开(公告)日：2025-03-25
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：US11156167B2
公开(公告)日：2021-10-26
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-
公开(公告)号：US10882633B2
公开(公告)日：2021-01-05
申请号：US16741005
申请日：2020-01-13
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 of: the turbine diameter at an axial location of the lowest pressure rotor stage a distance from a ground plane to the wing is in the range from 0.2 to 0.3.
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公开(公告)号：US20200346779A1
公开(公告)日：2020-11-05
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; 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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