公开(公告)号：US20200347748A1

公开(公告)日：2020-11-05

申请号：US16451537

申请日：2019-06-25

Applicant: ROLLS-ROYCE plc

Inventor： Chathura K. KANNANGARA ,  Jillian C. GASKELL ,  Stewart T. THORNTON ,  Timothy PHILP

IPC: F01D25/24 ,  F01D15/12

Abstract: A gas turbine engine for an aircraft and an engine core including: a compressor system, a first lower pressure compressor, a second, higher pressure compressor; an outer core casing surrounding the compressor system. 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, the first flange connection having a first flange radius, wherein the first flange connection is the 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, and a front mount arranged to be connected to a pylon. A front mount position ratio of: axial   distance   between   the   first   flange   connection and   the   front   mount first   flange   radius is equal to or less than 1.18.

公开(公告)号：US20200347731A1

公开(公告)日：2020-11-05

申请号：US16778795

申请日：2020-01-31

Applicant: ROLLS-ROYCE plc

Inventor： Chathura K. KANNANGARA ,  Jillian C. GASKELL ,  Stewart T. THORNTON ,  Timothy PHILP

IPC: F01D5/14 ,  F02C7/36 ,  F02K3/06 ,  F01D25/28 ,  F16H1/28 ,  F01D25/24

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing. The engine includes a front mount arranged for connection to a pylon; and a fan located upstream of the engine core. The outer core casing includes a first flange connection that: is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A front mount position ratio of: axial   distance   between   the   first   flange   connection and   the   front   mount first   flange   radius is equal to or less than 1.18.

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

IPC: F02C7/06 ,  F02K3/04 ,  F01D25/16

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.

公开(公告)号：US20200347730A1

公开(公告)日：2020-11-05

申请号：US16778617

申请日：2020-01-31

Applicant: ROLLS-ROYCE plc

Inventor： Chathura K. KANNANGARA ,  Jillian C. GASKELL ,  Stewart T. THORNTON ,  Timothy PHILP

IPC: F01D5/14 ,  F02C7/36 ,  F02K3/06 ,  F16H1/28 ,  F01D25/28 ,  F01D25/24

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and inner and outer core casings that define a core working gas flow path (A) therebetween, which has an outer radius that defines a gas path radius. The outer core casing includes a first flange connection that: has a first flange radius, is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A gas path ratio of: first   flange   radius gas   path   radius is equal to or greater than 1.10.

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

IPC: F02C7/06 ,  F01D25/16 ,  F02C7/36 ,  F02K3/04

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.

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

IPC: F02C3/113 ,  B64D27/26 ,  F01D15/12 ,  F01D25/16 ,  F02C7/06

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.

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

IPC: F02K3/04 ,  F02C7/06

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.

公开(公告)号：US20210189956A1

公开(公告)日：2021-06-24

申请号：US16796051

申请日：2020-02-20

Applicant: ROLLS-ROYCE plc

Inventor： Chathura K. KANNANGARA ,  Jillian C. GASKELL ,  Punitha KAMESH

IPC: F02C3/113 ,  B64D27/26 ,  F01D25/16 ,  F01D15/12

Abstract: A gas turbine engine for an aircraft has an engine core comprising turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine, and having turbine blades, and the compressor being the lowest pressure compressor of the engine; fan located upstream of the engine core; and 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, the three bearings having two rearward bearings located downstream of the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade, and/or, when the turbine comprises four sets of turbine blades, downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

公开(公告)号：US20200347742A1

公开(公告)日：2020-11-05

申请号：US16520740

申请日：2019-07-24

Applicant: ROLLS-ROYCE plc

Inventor： Chathura K. KANNANGARA ,  Jillian C. GASKELL ,  Stewart T. THORNTON ,  Timothy PHILP

IPC: F01D13/00 ,  F01D5/12 ,  F01D5/30

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a 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, wherein 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 nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

公开(公告)号：US20200347732A1

公开(公告)日：2020-11-05

申请号：US16520516

申请日：2019-07-24

Applicant: ROLLS-ROYCE PLC

Inventor： Chathura K KANNANGARA ,  Jillian C. GASKELL ,  Stewart T THORNTON ,  Timothy PHILP

IPC: F01D5/14 ,  F02C3/06 ,  F01D25/24

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; an inner core casing provided radially inwardly of the compressor blades of the compressor system; and an outer core casing surrounding the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween. 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, the first flange connection having a first flange radius, A gas path radius is defined as the outer radius of the core gas flow path at the axial position of the first flange connection, and a gas path ratio of: first   flange   radius gas   path   radius is equal to or greater than 1.10.