公开(公告)号：US10746102B2

公开(公告)日：2020-08-18

申请号：US15923183

申请日：2018-03-16

Applicant: ROLLS-ROYCE plc

Inventor： James M. Pointon ,  Stephen J. Bradbrook

IPC: F02C7/36 ,  F01D25/24 ,  F02K3/06 ,  F02C9/18

Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), and an outer core casing provided outwardly of compressor blades, the inner core casing and outer core casing defining a core working gas flow path (B) therebetween; a fan (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) spaced radially outwardly from the first outer core casing, and wherein at an axial plane (E) of an inlet to the high pressure compressor, the second outer core casing has an inner radius at least 1.4 times the inner radius of the first outer core casing.

公开(公告)号：US11448087B2

公开(公告)日：2022-09-20

申请号：US17004601

申请日：2020-08-27

Applicant: ROLLS-ROYCE plc

Inventor： Stephen J. Bradbrook

IPC: F01D19/00 ,  F01D9/04 ,  F01D17/00 ,  F02C9/26 ,  F01D17/14

Abstract: A gas turbine engine comprising: a combustor configured to initiate combustion; and a turbine comprising a stator vane ring defining a plurality of passageways between adjacent vanes; wherein at least one of the passageways is provided with a restrictor which defines a temporary gas washed surface for the stator vane ring and is configured to be ablated upon initiation of combustion to reveal an operational gas washed surface of the stator vane ring. A method of starting a gas turbine engine is also described.

公开(公告)号：US11047339B2

公开(公告)日：2021-06-29

申请号：US16103329

申请日：2018-08-14

Applicant: ROLLS-ROYCE plc

Inventor： James M. Pointon ,  Stephen J. Bradbrook

IPC: F02K3/068 ,  F02K3/06 ,  F02C7/04 ,  F02C3/107 ,  F02C7/36 ,  F04D29/54 ,  B64D33/02 ,  F01D1/02

Abstract: An aircraft gas turbine engine comprises a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow (B) and a core flow (A) in use. The engine includes a reduction gearbox which couples the fan to the fan drive turbine and a core compressor arrangement. The core compressor arrangement has a core inlet at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls, and at least a first compressor rotor blade provided at an upstream end of the compressor arrangement. The radially inner wall of the core inlet defines a first diameter (DINLET), and a root leading edge of the first compressor rotor blade defines a second diameter (DCOMP). A first ratio (DINLET:DCOMP) of the first diameter (DCOMP) to the second diameter (DCOMP) is greater than or equal to 1.4.