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公开(公告)号：US20240011433A1
公开(公告)日：2024-01-11
申请号：US18215868
申请日：2023-06-29
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J. BRADBROOK , William J. SMITH , Jonathan P. BRADLEY
IPC: F02C3/06
CPC classification number: F02C3/06 , F05D2220/36
Abstract: A novel configuration for axial flow gas turbine engine for aircraft has an engine core having a core length and including first turbine, axial compressor, and drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.
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公开(公告)号：US20220112817A1
公开(公告)日：2022-04-14
申请号：US17495418
申请日：2021-10-06
Applicant: ROLLS-ROYCE PLC
Inventor： Stephen J. BRADBROOK , Martin N. GOODHAND , Paul M. HIELD , Andrew PARSLEY , Natalie C. WONG , Robert J. CORIN , Thomas S. BINNINGTON
Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.
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公开(公告)号：US20230167768A1
公开(公告)日：2023-06-01
申请号：US17940055
申请日：2022-09-08
Applicant: ROLLS-ROYCE plc
Inventor： Natalie C. WONG , Jonathan A. CHERRY , Paul R. DAVIES , David A. JONES , Andrew J. NEWMAN , Benjamin J. SELLERS , Stephen J. BRADBROOK
CPC classification number: F02C7/12 , F02C6/20 , F02C7/32 , F05D2220/323 , F05D2220/76 , F05D2260/213
Abstract: An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.
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公开(公告)号：US20170369179A1
公开(公告)日：2017-12-28
申请号：US15605125
申请日：2017-05-25
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J. BRADBROOK
CPC classification number: B64D35/04 , B64D27/12 , B64D27/18 , B64D2027/005 , F02C3/04 , F02C3/107 , F02C3/145 , F02C6/02 , F02C7/18 , F02C7/36 , F05D2220/323 , F05D2240/35 , F05D2260/213
Abstract: An aircraft gas turbine engine (110) comprises first and second non-coaxial propulsors (113a, 113b), each propulsor (113a, 113b) being driven by a common gas turbine engine core (176) comprising a propulsor drive turbine (143) arranged to drive the first and second propulsors (113a, 113b) via a propulsor drive coupling (127). The core (176) further comprises a first core module (190) comprising a first compressor (129) and a first turbine (131) interconnected by a first shaft (177), and a second core module (191) comprising a second compressor (128) and the propulsor drive turbine (143) interconnected by a second shaft (127), the first and second core modules (190, 191) being axially spaced.
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公开(公告)号：US20230167775A1
公开(公告)日：2023-06-01
申请号：US17940419
申请日：2022-09-08
Applicant: ROLLS-ROYCE PLC
Inventor： Benjamin J. SELLERS , Andrew J. NEWMAN , Gordon MARGARY , Paul R. DAVIES , Stephen J. BRADBROOK
CPC classification number: F02C7/32 , F02C6/00 , F02C6/20 , F02C7/12 , F05D2220/76 , F05D2220/323 , F05D2260/213
Abstract: A gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a total electrical power PEM1 (W), and the gas turbine engine is configured to generate a total shaft power PSHAFT (W); and a ratio R of:
R
=
Total Shaft Power =
P
S
H
A
F
T
Total Electrical Power Generated = P
E
M
1
is in a range of between 0.005 and 0.020.-
Patent Agency Ranking
公开(公告)号：US20220112845A1
公开(公告)日：2022-04-14
申请号：US17484371
申请日：2021-09-24
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J. BRADBROOK , Martin N. GOODHAND , Paul M. HIELD , Andrew PARSLEY , Natalie C. WONG , Robert J. CORIN , Thomas S. BINNINGTON
Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.
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公开(公告)号：US20220112840A1
公开(公告)日：2022-04-14
申请号：US17494523
申请日：2021-10-05
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J. BRADBROOK , Martin N. GOODHAND , Paul M. HIELD , Andrew PARSLEY , Natalie C. WONG , Robert J. CORIN , Thomas S. BINNINGTON
Abstract: A turbofan gas turbine engine includes heat exchanger module, fan assembly, compressor, turbine and exhaust modules. The fan includes a plurality of fan blades. The heat exchanger in fluid communicates with the fan assembly by an inlet duct, and the heat exchanger includes a plurality of radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger is divided between a set of vane airflows and a set of channel airflows. Each vane airflow has a vane mass flow rate FlowVane, and each channel air flow has a channel mass flow rate FlowChan. Each hollow vane includes, an inlet, heat transfer, and exhaust portions, with the inlet portion comprising a diffuser element and the heat transfer portion including at least one heat transfer element. The diffuser element causes FlowVane to be lower than FlowChan.
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公开(公告)号：US20190048826A1
公开(公告)日：2019-02-14
申请号：US16103329
申请日：2018-08-14
Applicant: ROLLS-ROYCE plc
Inventor： James M. POINTON , Stephen J. BRADBROOK
Abstract: An aircraft gas turbine engine comprises a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow (B) and a core flow (A) in use. The engine includes a reduction gearbox which couples the fan to the fan drive turbine and a core compressor arrangement. The core compressor arrangement has a core inlet at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls, and at least a first compressor rotor blade provided at an upstream end of the compressor arrangement. The radially inner wall of the core inlet defines a first diameter (DINLET), and a root leading edge of the first compressor rotor blade defines a second diameter (DCOMP). A first ratio (DINLET:DCOMP) of the first diameter (DCOMP) to the second diameter (DCOMP) is greater than or equal to 1.4.
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公开(公告)号：US20210079807A1
公开(公告)日：2021-03-18
申请号：US17004601
申请日：2020-08-27
Applicant: ROLLS-ROYCE plc
Inventor： Stephen J. BRADBROOK
Abstract: A gas turbine engine comprising: a combustor configured to initiate combustion; and a turbine comprising a stator vane ring defining a plurality of passageways between adjacent vanes; wherein at least one of the passageways is provided with a restrictor which defines a temporary gas washed surface for the stator vane ring and is configured to be ablated upon initiation of combustion to reveal an operational gas washed surface of the stator vane ring. A method of starting a gas turbine engine is also described.
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公开(公告)号：US20180283282A1
公开(公告)日：2018-10-04
申请号：US15923183
申请日：2018-03-16
Applicant: ROLLS-ROYCE plc
Inventor： James M. POINTON , Stephen J. BRADBROOK
CPC classification number: F02C7/36 , F01D25/24 , F02C9/18 , F02K3/06 , F05D2220/3216 , F05D2220/323 , F05D2240/60 , Y02T50/671
Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), and an outer core casing provided outwardly of compressor blades, the inner core casing and outer core casing defining a core working gas flow path (B) therebetween; a fan (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) spaced radially outwardly from the first outer core casing, and wherein at an axial plane (E) of an inlet to the high pressure compressor, the second outer core casing has an inner radius at least 1.4 times the inner radius of the first outer core casing.
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