公开(公告)号：US20240102431A1

公开(公告)日：2024-03-28

申请号：US18215864

申请日：2023-06-29

Applicant: ROLLS-ROYCE plc

Inventor： Stephen J BRADBROOK ,  William J SMITH ,  Jonathan P BRADLEY

IPC: F02K3/06 ,  F01D9/04 ,  F01D25/24

CPC classification number: F02K3/06 ,  F01D9/041 ,  F01D25/24 ,  F05D2240/35

Abstract: A novel configuration for an axial flow gas turbine engine for aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a 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 combustor volume (in litres) which when divided by fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit an electric motor to be located within fan hub diameter.

公开(公告)号：US20240011432A1

公开(公告)日：2024-01-11

申请号：US18215860

申请日：2023-06-29

Applicant: ROLLS-ROYCE plc

Inventor： Stephen J BRADBROOK ,  William J SMITH ,  Jonathan P BRADLEY

IPC: F02C3/04

CPC classification number: F02C3/04 ,  F05D2220/36

Abstract: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a 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 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 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 multistage fan configured to permit electric motor to be located within fan hub diameter.

公开(公告)号：US20210010384A1

公开(公告)日：2021-01-14

申请号：US16915032

申请日：2020-06-29

Applicant: ROLLS-ROYCE plc

Inventor： Jonathan P BRADLEY

IPC: F01D15/10 ,  F01D15/12

Abstract: An aircraft gas turbine engine (10) comprises a main engine shaft (22) arranged to couple a turbine (17) and a compressor (13), the main engine shaft (22) defining an axial direction (9). The gas turbine engine (10) further comprises at least one radially extending offtake shaft (27) coupled to the main engine shaft (22), and a radially extending electric machine (25a, 25b) coupled to the radially extending offtake shaft (22).