公开(公告)号：US10577956B2

公开(公告)日：2020-03-03

申请号：US15907538

申请日：2018-02-28

Applicant: ROLLS-ROYCE plc

Inventor： Richard Northall ,  Anthony J. Rae ,  Michael S. Krautheim ,  Alastair D. Walker ,  Jonathan F. Carrotte ,  Ian Mariah

IPC: F01D5/14 ,  F01D9/04 ,  F01D9/06 ,  F01D9/02 ,  F02K3/04

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The space to chord ratio of the aerodynamic vanes that are closest to the suction surface of a strut-vane is higher than the space to chord ratio of aerodynamic vanes that are closest to a pressure surface of the strut-vane. The arrangement allows the duct 100 to be axially short.

公开(公告)号：US10697471B2

公开(公告)日：2020-06-30

申请号：US15907531

申请日：2018-02-28

Applicant: ROLLS-ROYCE plc

Inventor： Richard Northall ,  Anthony J. Rae ,  Michael S. Krautheim ,  Alastair D. Walker ,  Jonathan F. Carrotte ,  Ian Mariah

IPC: F04D29/54 ,  F01D25/16 ,  F01D9/04 ,  F02C3/06 ,  F01D5/14

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The arrangement allows the duct 100 to be axially short.

公开(公告)号：US20190338707A1

公开(公告)日：2019-11-07

申请号：US16386542

申请日：2019-04-17

Applicant: Rolls-Royce plc

Inventor： Peter A. Beecroft ,  Gareth M. Armstrong ,  Bharat R. Koli ,  Alastair D. Walker

IPC: F02C7/18

Abstract: The present disclosure relates to a compressor assembly for a gas turbine engine. The compressor assembly comprises a compressor having a compressor stage with a plurality of rotor blades mounted on the rim of a rotor disc within a core air flow path. There is a purge channel extending from a radially inner purge gas source to at least one radially outer outlet at the rim of the compressor stage. The purge channel comprises at least one swirl element and/or at least one orifice for swirling/deflecting a purge gas flowing within the purge channel. In use, a flow of purge gas exits the outlet(s) at least partly in the direction of the core air flow path.