-
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.
-
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.
-
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.
-
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.
-
Patent Agency Ranking
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.
-
公开(公告)号：US20210115797A1
公开(公告)日：2021-04-22
申请号：US17082886
申请日：2020-10-28
Applicant: ROLLS-ROYCE plc
Inventor： Chathura K KANNANGARA , Jillian C GASKELL , Stewart T THORNTON , Timothy PHILP
Abstract: A gas turbine engine includes an engine core and a fan located upstream of the engine core. The engine core includes: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. The first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor. A fan diameter ratio of: first flange radius fan diameter is equal to or greater than 0.125.
-
公开(公告)号：US20200347749A1
公开(公告)日：2020-11-05
申请号：US16505816
申请日：2019-07-09
Applicant: ROLLS-ROYCE PLC
Inventor： Chathura K KANNANGARA , Jillian C GASKELL , Stewart T THORNTON , Timothy PHILP
Abstract: A gas turbine engine for an aircraft includes: an engine core with: a compressor system including a first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system. The gas turbine engine further includes a fan located upstream of the engine core with a plurality of fan blades and having a fan diameter. The outer core casing includes a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection which has a first flange radius, wherein the first flange connection is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor.
-
公开(公告)号：US20220098984A1
公开(公告)日：2022-03-31
申请号：US17381767
申请日：2021-07-21
Applicant: ROLLS-ROYCE plc
Inventor： Chathura K KANNANGARA , Jillian C GASKELL , Stewart T THORNTON , Timothy PHILP
Abstract: A gas turbine engine includes an engine core, a fan located upstream of the engine core, a nacelle surrounding the engine core and defining a bypass duct, and a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle. The engine core includes a compressor system, and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. An axial midpoint of a radially inner edge of the fan OGV is defined as the fan OGV root centrepoint. A fan OGV root position to fan diameter ratio of: an axial distance between the first flange connection and the fan OGV root centrepoint the fan diameter is equal to or less than 0.33.
-
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.
-
-
-
-
-
-
-
-
Search Results
9
records
(took 0.012 seconds)
