公开(公告)号：US11339717B2

公开(公告)日：2022-05-24

申请号：US17113910

申请日：2020-12-07

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Amarveer S Mann

IPC: B64D13/08 ,  F02C6/08 ,  F02C7/18 ,  B64D13/06 ,  F02C7/057 ,  F02C7/04 ,  F02C7/36

Abstract: The present disclosure relates to a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft; and an environmental control system mounted on the engine core comprising a first air passage arranged to deliver air from outside the engine core to an aircraft cabin and/or for wing anti icing, a subsidiary compressor located in the first air passage and arranged to compress air in the first air passage, the subsidiary compressor being powered by the core shaft, and a second air passage arranged to inject air from the compressor into the first air passage.

公开(公告)号：US10961916B2

公开(公告)日：2021-03-30

申请号：US16443938

申请日：2019-06-18

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F02C7/36 ,  F02C3/113 ,  F01D15/12

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.