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公开(公告)号：US20230242260A1
公开(公告)日：2023-08-03
申请号：US18155351
申请日：2023-01-17
Applicant: ROLLS-ROYCE plc , Rolls-Royce Corporation
Inventor： Christopher A. MURRAY , Nicholas HOWARTH , Daniel SWAIN , Ian J. BOUSFIELD
CPC classification number: B64D13/06 , F04D25/04 , F02C7/32 , B64D2013/0603
Abstract: There is provided a blower system for providing air to an airframe system, comprising a rotor configured to be mechanically coupled to a spool 440 of a gas turbine engine, wherein the rotor is configured to: in a blower mode, be driven to rotate by the spool to discharge air to an airframe discharge port for supply to an airframe system; and, in an engine drive mode, receive air from an external air source via an impingement port that is configured to direct the received air onto the rotor and thereby drive the rotor to rotate to drive the spool to.
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公开(公告)号：US20140017067A1
公开(公告)日：2014-01-16
申请号：US13914882
申请日：2013-06-11
Applicant: ROLLS-ROYCE PLC
Inventor： Richard Geoffrey STRETTON , Nicholas HOWARTH
IPC: B64C11/00
CPC classification number: B64C11/001 , B64D27/00 , B64D2027/005 , F02C3/067 , F02C7/04 , F02C7/05 , F02K3/025 , F02K3/04 , F05D2220/323 , F05D2220/325 , F05D2240/302 , F05D2250/324 , F05D2250/51 , Y02T50/66 , Y02T50/671 , Y02T50/673
Abstract: A gas turbine engine (10) having an axial flow direction (X) therethrough in use. The gas turbine engine (10) comprises one or more rotor stages each comprising at least one rotor blade (120) having a root portion (122). The gas turbine engine (10) comprises a shroud (122) located upstream of one or more of the rotor stages relative to the axial flow direction (X). The shroud (122) defines a through passageway (128) extending between an inlet (130) and an outlet (132) which comprises a diffuser region (138). The diffuser region (138) is configured to reduce the axial velocity of air exiting the outlet (132) relative to air entering the diffuser portion (138) in use, wherein the outlet (132) is located such that air exiting the outlet (132) is directed substantially to the root portion (122) only of the rotor blades (120).
Abstract translation: 一种在使用中具有轴流动方向（X）的燃气轮机（10）。 燃气涡轮发动机（10）包括一个或多个转子级，每个转子级包括至少一个具有根部（122）的转子叶片（120）。 燃气涡轮发动机（10）包括相对于轴向流动方向（X）定位在一个或多个转子级的上游的护罩（122）。 护罩（122）限定在入口（130）和包括扩散器区域（138）的出口（132）之间延伸的通道（128）。 扩散器区域（138）构造成在使用时相对于进入扩散器部分（138）的空气减少离开出口（132）的空气的轴向速度，其中出口（132）定位成使得离开出口（132）的空气 ）仅基本上指向转子叶片（120）的根部（122）。
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公开(公告)号：US20240317406A1
公开(公告)日：2024-09-26
申请号：US18609241
申请日：2024-03-19
Applicant: Rolls-Royce plc
Inventor： Christopher A. MURRAY , Nicholas HOWARTH
CPC classification number: B64D13/04 , F01D11/04 , F02C7/06 , F02C7/28 , F05D2240/55 , F05D2260/98
Abstract: A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; an oil-lubricated bearing for supporting rotation of the rotor, wherein the oil-lubricated bearing is disposed within a bearing chamber; and a seal assembly for restricting oil from the bearing chamber from reaching an inlet of the blower compressor, wherein the seal assembly comprises: a seal disposed between the bearing chamber and an air buffer chamber; an air buffer inlet configured to receive a flow of pressurized air into the air buffer chamber; and an air buffer outlet configured to allow the flow of pressurized air to be exhausted from the air buffer chamber together with oil from the bearing chamber that has passed through the seal and become entrained within the flow of pressurized air.
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公开(公告)号：US20230079630A1
公开(公告)日：2023-03-16
申请号：US17987516
申请日：2022-11-15
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas HOWARTH , Gareth M. ARMSTRONG
Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
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公开(公告)号：US20220119120A1
公开(公告)日：2022-04-21
申请号：US17562609
申请日：2021-12-27
Applicant: ROLLS-ROYCE PLC
Inventor： Gareth M ARMSTRONG , Nicholas HOWARTH
Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
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Patent Agency Ranking
公开(公告)号：US20210172374A1
公开(公告)日：2021-06-10
申请号：US17113910
申请日：2020-12-07
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas HOWARTH , Amarveer S. MANN
Abstract: The present disclosure relates to a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft; and an environmental control system mounted on the engine core comprising a first air passage arranged to deliver air from outside the engine core to an aircraft cabin and/or for wing anti icing, a subsidiary compressor located in the first air passage and arranged to compress air in the first air passage, the subsidiary compressor being powered by the core shaft, and a second air passage arranged to inject air from the compressor into the first air passage.
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公开(公告)号：US20240102399A1
公开(公告)日：2024-03-28
申请号：US18463633
申请日：2023-09-08
Applicant: Rolls-Royce plc
Inventor： Christopher A. MURRAY , Nicholas HOWARTH
CPC classification number: F01D11/001 , F01D15/08 , F05D2240/56
Abstract: There is provided a dynamic sealing assembly for a rotary machine, comprising a primary sandwich plate, a secondary sandwich plate and a bristle pack. The primary sandwich plate comprises a plurality of primary vane openings, and the secondary sandwich plate comprises a plurality of secondary vane openings. The bristle pack comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings overlies and aligns with a respective secondary vane opening to form a vane channel for receiving a vane along a longitudinal axis of the dynamic sealing assembly. The bristle pack is configured to: provide a brush seal between each vane received within the respective vane channels and the dynamic sealing assembly; and allow relative movement between the dynamic sealing assembly and the vane received within each vane channel along the longitudinal axis.
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公开(公告)号：US20230242264A1
公开(公告)日：2023-08-03
申请号：US18098463
申请日：2023-01-18
Applicant: ROLLS-ROYCE PLC
Inventor： Gareth M. ARMSTRONG , Nicholas HOWARTH
CPC classification number: B64D27/10 , F01D5/28 , F02C3/04 , F02C3/06 , F02C3/107 , F02C7/04 , F02K3/06 , F02K3/068 , B64D2033/0286 , F05D2220/36
Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
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公开(公告)号：US20200291865A1
公开(公告)日：2020-09-17
申请号：US16443938
申请日：2019-06-18
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas HOWARTH , Gareth M ARMSTRONG
Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
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公开(公告)号：US20200291785A1
公开(公告)日：2020-09-17
申请号：US16713666
申请日：2019-12-13
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas HOWARTH , Gareth M ARMSTRONG
IPC: F01D5/14 , F02C3/02 , F04D19/02 , F02C7/36 , F02K3/068 , F02C3/04 , F02C3/107 , F01D5/28 , F04D29/68 , F02K3/06 , F02C7/04
Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
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