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公开(公告)号：US20230349327A1
公开(公告)日：2023-11-02
申请号：US18213392
申请日：2023-06-23
Applicant: ROLLS-ROYCE PLC
Inventor： Jillian C GASKELL , Chathura K KANNANGARA , Punitha KAMESH
CPC classification number: F02C7/06 , F01D25/16 , F02K3/04 , F05D2240/50 , F02C7/36 , F05D2200/14 , F05D2220/323 , F05D2240/30
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.
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公开(公告)号：US20210189963A1
公开(公告)日：2021-06-24
申请号：US16824270
申请日：2020-03-19
Applicant: ROLLS-ROYCE plc
Inventor： Chathura K KANNANGARA , Punitha KAMESH , Jillian C GASKELL
Abstract: An aircraft gas turbine engine has an engine core having a turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine and having a turbine length being the distance between the roots of the most upstream turbine blade at its leading edge and of the most downstream turbine blade trailing edge, and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine; and a gearbox receiving an input from the core shaft and outputting drive to the fan. The engine core has three bearings to support the core shaft, the three bearings having a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings, and wherein further a minor span to turbine length ratio of: minor span tu rbine length is equal to or less than 1.05.
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公开(公告)号：US20210189908A1
公开(公告)日：2021-06-24
申请号：US16796055
申请日：2020-02-20
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C GASKELL , Chathura K KANNANGARA , Punitha KAMESH
Abstract: A gas turbine engine for an aircraft has an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, and two rearward bearings, and wherein the forward most rearward bearing has a bearing stiffness defined by the radial displacement caused by the application of a radial force at the axial centerpoint of the bearing, and wherein a stiffness ratio of the bearing stiffness at the forward most rearward bearing to the minor span is in the range from 0.08 to 0.5 kN/mm2.
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公开(公告)号：US20210190009A1
公开(公告)日：2021-06-24
申请号：US16823520
申请日：2020-03-19
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C GASKELL , Chathura K KANNANGARA , Punitha KAMESH
Abstract: An engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, and wherein the forwardmost rearward bearing has a bearing stiffness in the range of 30 kN/mm to 100 kN/mm, the bearing stiffness being defined by the radial displacement caused by the application of a radial force at the axial centrepoint of the bearing.
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公开(公告)号：US20210189971A1
公开(公告)日：2021-06-24
申请号：US16809772
申请日：2020-03-05
Applicant: ROLLS-ROYCE PLC
Inventor： Jillian C GASKELL , Chathura K KANNANGARA , Punitha KAMESH
Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting the two, the turbine and compressor being the lowest pressure turbine and compressor of the engine, the core shaft having a running speed range between 1500 rpm and 6200 rpm; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The engine core includes a forward bearing and two rearward bearings arranged to support the core shaft, and the core shaft having a length between the forward bearing and the rearmost rearward bearing and a minor span between the rearward bearings, and the length ratio of the minor span to the core shaft length is equal to or less than 0.14.
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Patent Agency Ranking
公开(公告)号：US20220235702A1
公开(公告)日：2022-07-28
申请号：US17718923
申请日：2022-04-12
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C. GASKELL , Chathura K. KANNANGARA , Punitha KAMESH
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan including a plurality of fan blades located upstream of the engine core. The fan has a fan diameter in the range from 240 cm to 280 cm. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine includes a total of three sets of turbine blades. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the blade.
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公开(公告)号：US20210189962A1
公开(公告)日：2021-06-24
申请号：US16809831
申请日：2020-03-05
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C GASKELL , Chathura K KANNANGARA , Punitha KAMESH
Abstract: An engine core including a turbine, compressor, and core shaft connecting the turbine and compressor, the turbine being the lowest pressure turbine, the core shaft having a running speed range from 1500-6200 rpm, and the compressor being the lowest pressure compressor; a fan located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, the core shaft having a length between the forward and the rearmost rearward bearing ranging from 1800-2900 mm, and a minor span between two rearward bearings ranging from 250-350 mm, wherein there is no primary resonance of the core shaft between the forward and forwardmost rearward bearing within the running speed range of the core shaft.
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公开(公告)号：US20240352891A1
公开(公告)日：2024-10-24
申请号：US18749076
申请日：2024-06-20
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C. GASKELL , Chathura K. KANNANGARA , Punitha KAMESH
CPC classification number: F02C7/06 , F01D25/16 , F02K3/04 , F02C7/36 , F05D2200/14 , F05D2220/323 , F05D2240/30 , F05D2240/50
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.
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公开(公告)号：US20220136434A1
公开(公告)日：2022-05-05
申请号：US17472146
申请日：2021-09-10
Applicant: ROLLS-ROYCE plc
Inventor： Chathura K. KANNANGARA , Jillian C. GASKELL , Punitha KAMESH
Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core. The engine core further includes three bearings arranged to support the core shaft including a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings. A minor span to turbine length ratio is equal to or less than 1.05.
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公开(公告)号：US20210190008A1
公开(公告)日：2021-06-24
申请号：US16809984
申请日：2020-03-05
Applicant: ROLLS-ROYCE plc
Inventor： Jillian C. GASKELL , Chathura K. KANNANGARA , Punitha KAMESH
Abstract: An aircraft gas turbine engine has an engine core with a turbine, compressor, and core shaft connecting the turbine to the compressor, a fan upstream of the engine core; and a gearbox. The engine core has three bearings, one forward, two rearward, to support the core shaft, a minor span being the axial distance between the two rearward bearings. A first blade to bearing ratio of the minor span divided by the product of the mass, radius at mid-height, and the square of the angular velocity at cruise for a blade of the lowest pressure set may have a value in the range from 2.0×10−6 to 7.5×10−6 kg−1.rad−2.s2. A second blade to bearing ratio of the minor span divided by the product of mass and radius at mid-height for a blade of the lowest pressure set may have a value in the range from 0.8 to 6.0 kg−1.
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