公开(公告)号：EP3981968A1

公开(公告)日：2022-04-13

申请号：EP21198481.0

申请日：2021-09-23

申请人： Rolls-Royce plc

发明人： Bradbrook, Stephen J ,  Goodhand, Martin N ,  Hield, Paul M ,  Parsley, Andrew ,  Wong, Natalie C ,  Corin, Robert J ,  Binnington, Thomas S

IPC分类号： F02C7/04 ,  F02C7/047 ,  F02C7/14 ,  F28D1/02 ,  F01D9/06 ,  F01D25/02 ,  F02K3/072 ,  F01D9/04

摘要： A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly comprises a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module comprises a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels. Each vane airflow having a vane mass flow rate Flow Vane , and each channel air flow has a channel mass flow rate Flow Chan . At least one of the hollow vanes accommodates at least one heat transfer element for the transfer of heat from a first fluid contained within the or each heat transfer element to the or each corresponding vane airflow passing over a surface of the or each heat transfer element.
In use, a Vane Airflow Ratio parameter V AR is defined as: VAR = Flow VaneTot Flow ChanTot where:
Flow VaneTot = total mass flow rate of the vane mass flow rates, Flow Vane ; and
Flow ChanTot = total mass flow rate of the channel mass flow rates, Flow Chan ; and the V AR parameter is in the range of 0.05 to 3.0.

公开(公告)号：EP3444468A1

公开(公告)日：2019-02-20

申请号：EP18183848.3

申请日：2018-07-17

申请人： Rolls-Royce plc

发明人： Pointon, James M ,  Bradbrook, Stephen J

IPC分类号： F02K3/06 ,  F04D29/54

摘要： An aircraft gas turbine engine (10) comprises a fan (16) coupled to a fan drive turbine (26), the fan (14) being configured to provide a bypass flow (B) and a core flow (A) in use. The engine (10) includes a reduction gearbox (36) which couples the fan (16) to the fan drive turbine (26) and a core compressor arrangement (18, 20). The core compressor arrangement (18, 20) has a core inlet (38) at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls (72, 74), and at least a first compressor rotor blade (40) provided at an upstream end of the compressor arrangement (18, 20). The radially inner wall (72) of the core inlet (38) defines a first diameter (D INLET ), and a root leading edge (84) of the first compressor rotor blade (40) defines a second diameter (D COMP ). A first ratio (D INLET :D COMP ) of the first diameter (D COMP ) to the second diameter (D COMP ) is greater than or equal to 1.4.