公开(公告)号：US20210164356A1

公开(公告)日：2021-06-03

申请号：US17084821

申请日：2020-10-30

Applicant: ROLLS-ROYCE plc

Inventor： Fernando L. TEJERO EMBUENA ,  David G. MACMANUS ,  Christopher TJ SHEAF

IPC: F01D9/04 ,  F01D25/24 ,  F02C7/04

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (θscarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (θdroop) relative to the longitudinal centre line. A relationship between the droop angle (θdroop) and the scarf angle (θscarf) is given by: θdroop=θscarf/1.5±1 degree.

公开(公告)号：US20210062760A1

公开(公告)日：2021-03-04

申请号：US16985865

申请日：2020-08-05

Applicant: ROLLS-ROYCE plc

Inventor： Christopher TJ SHEAF ,  Chia Hui LIM

IPC: F02K3/02 ,  F02C3/113 ,  F01D25/34

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine and compressor, connected by a core shaft. The fan is upstream of the core and includes a plurality of fan blades, and has a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60. The nacelle at least partially surrounds the core and fan. The gearbox receives input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The engine has a bypass ratio greater than 10. The engine has a local diffuser angle from 0 to 18 degrees, a local peak diffuser angle from 0 to 22 degrees, and/or a bulk diffuser angle from 0 to 15 degrees.

公开(公告)号：US20210381397A1

公开(公告)日：2021-12-09

申请号：US17320297

申请日：2021-05-14

Applicant: ROLLS-ROYCE plc

Inventor： Fernando L. TEJERO EMBUENA ,  David G. MACMANUS ,  Christopher TJ SHEAF

IPC: F01D25/24 ,  B64D27/12 ,  B64D29/02

Abstract: A nacelle for a gas turbine engine includes a leading edge at an upstream side of the nacelle. The nacelle further includes a trailing edge at a downstream side of the nacelle. The nacelle further includes an outer surface disposed between the leading edge and the trailing edge. The nacelle further includes a concave section continuous with the outer surface and disposed proximal to the trailing edge. The concave section includes an upstream end and a downstream end spaced apart from the upstream end. The concave section is curved radially inwards relative to the outer surface between the upstream end and the downstream end.

公开(公告)号：US20210062762A1

公开(公告)日：2021-03-04

申请号：US16985912

申请日：2020-08-05

Applicant: ROLLS-ROYCE plc

Inventor： Christopher TJ SHEAF ,  Chia Hui LIM

IPC: F02K3/06 ,  F02C7/36 ,  F02C3/06

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is upstream of the core and includes a plurality of fan blades, and has a diameter greater than 2.0 m. The air intake is upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The nacelle at least partially surrounds the core and fan. The gearbox receives input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft. The engine has a bypass ratio greater than 10. The nacelle has a length and the ratio of the length of the nacelle to the fan diameter is 0.4 to 2.5.

公开(公告)号：US20210062759A1

公开(公告)日：2021-03-04

申请号：US16985817

申请日：2020-08-05

Applicant: ROLLS-ROYCE plc

Inventor： Chia Hui LIM ,  Christopher TJ SHEAF

IPC: F02K3/02 ,  F01D25/34 ,  F02C3/113

Abstract: A gas turbine engine for an aircraft includes an engine core, a fan, an air intake and a gearbox. The engine core includes a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan is upstream of the engine core and includes a plurality of fan blades, the fan having a diameter greater than 2.0 m. The air intake is located upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60. The gearbox receives an input from the core shaft and outputs drive to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has a bypass ratio greater than 10; and the air intake defines a highlight area, a throat area and a diffuser area, wherein the ratio of the throat area to fan face area is from 0.94 to 1.05.