公开(公告)号：US20190382122A1

公开(公告)日：2019-12-19

申请号：US16416708

申请日：2019-05-20

Applicant: ROLLS-ROYCE plc

Inventor： Christopher T J SHEAF ,  Richard G. STRETTON ,  Chia Hui LIM

IPC: B64D27/18 ,  F02C7/04 ,  B64D33/02

Abstract: A gas turbine engine comprises a pylon attachment, a shaft defining an engine centreline which lies in an engine central plane intersecting the pylon attachment, a fan defining a fan plane normal to the engine centreline and an intake upstream of the fan plane. The geometric centreline of the intake coincides with the engine centreline at an axial position corresponding to the downstream end of the intake and curves away from the engine centreline upstream of said axial position. The engine may be mounted on one side of an aircraft such that the orientation of the highlight plane of the intake is aligned to the air flow field of the aircraft on that side during flight.

公开(公告)号：US20210355873A1

公开(公告)日：2021-11-18

申请号：US17231130

申请日：2021-04-15

Applicant: ROLLS-ROYCE plc

Inventor： Robert E CHRISTIE ,  David G MACMANUS ,  Christopher T J SHEAF

IPC: F02C7/045 ,  F02K3/06 ,  B64D33/02

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The nacelle further includes a protrusion extending radially inward from the air intake downstream of the intake lip. The protrusion extends circumferentially by a protrusion angle (θp) with respect to the longitudinal centre line of the gas turbine engine.

公开(公告)号：US20200023984A1

公开(公告)日：2020-01-23

申请号：US16460307

申请日：2019-07-02

Applicant: ROLLS-ROYCE PLC

Inventor： Chia Hui LIM ,  Richard G STRETTON ,  Christopher T J SHEAF

IPC: B64D27/26 ,  B64D29/06

Abstract: A mounting arrangement for mounting an aircraft gas turbine engine to an aircraft includes an engine nacelle with a distal assembly including a part annular engine cowl, a gas turbine engine core housing surrounded by the cowl and a distal bifurcation extending between the engine core housing and engine cowl in a first direction to define a first axis. The mounting arrangement includes a proximal assembly having a mount configured to mount the proximal assembly to the engine core housing. The proximal assembly includes a pylon configured to mount the proximal assembly to mounting location such as a wing of the aircraft at an engine mounting location. The pylon extends in a line between the wing and the engine core housing to define a second axis which is normal to a distal surface of the wing at the engine mounting location and is non-parallel to the vertical axis.

公开(公告)号：US20190383215A1

公开(公告)日：2019-12-19

申请号：US16416766

申请日：2019-05-20

Applicant: ROLLS-ROYCE plc

Inventor： Christopher T J SHEAF ,  Richard G STRETTON ,  Chia Hui LIM

IPC: F02C7/04

Abstract: A gas turbine engine (100) for an aircraft comprises a pylon attachment (112) and a shaft (108) defining an engine centreline (110). The engine centreline lies in an engine central plane (120) which intersects the pylon attachment. The gas turbine engine comprises an intake (104) having a non-axisymmetric geometry and a medial plane (130) defining left and right halves of the intake. The left and right halves are configured for at least one of optimum cross wind performance, optimum incidence performance and optimum cruise performance when the medial plane is aligned with a vertical plane. The intake is installed so that the medial plane is angularly offset with respect to the engine central plane. The engine may be installed on a wing of an aircraft with the medial plane closer to its optimal orientation than is the case for a conventional engine.

公开(公告)号：US20170328280A1

公开(公告)日：2017-11-16

申请号：US15494665

申请日：2017-04-24

Applicant: ROLLS-ROYCE plc

Inventor： Zahid M HUSSAIN ,  Christopher T J SHEAF ,  Mark J WILSON ,  David G MACMANUS

IPC: F02C7/141 ,  F28F3/02 ,  F28F27/02 ,  F28F13/12

CPC classification number: F02C7/141 ,  F02C7/185 ,  F02K3/115 ,  F05D2260/22141 ,  F28F3/02 ,  F28F13/12 ,  F28F27/02 ,  F28F2250/06 ,  Y02T50/671 ,  Y02T50/675

Abstract: A heat sink for location in a fluid flow, including a heat sink base and a plurality of heat dissipating elements, such as elongate fins, extending from the surface of the heat sink base. In certain arrangements the heat sink is provided with a diversion flow passageway for diverting a fraction of fluid flow away from the heat dissipating elements. In other arrangements there may be two arrays of elongate fins laterally offset. In yet a further arrangement the heat sink may be configured to promote the generation of at least one vortex.

公开(公告)号：US20210355872A1

公开(公告)日：2021-11-18

申请号：US17231421

申请日：2021-04-15

Applicant: ROLLS-ROYCE PLC

Inventor： Robert E. CHRISTIE ,  David G. MACMANUS ,  Christopher T J SHEAF

IPC: F02C7/04

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The diffuser further includes a diffuser angle (θdiff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (θdiff) is from about 0 degrees to about 12 degrees.

公开(公告)号：US20210254507A1

公开(公告)日：2021-08-19

申请号：US17152075

申请日：2021-01-19

Applicant: ROLLS-ROYCE plc

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

IPC: F01D25/24 ,  B64D27/12 ,  F02C7/04

Abstract: A nacelle for housing a fan within a gas turbine engine having a longitudinal centre line includes a leading edge and a trailing edge. A nacelle length (Lnac) is defined as an axial distance between the leading edge and the trailing edge. An azimuthal angle (θ) is defined about the longitudinal centre line. The nacelle length (Lnac) varies azimuthally. The nacelle length (Lnac) decreases azimuthally from an inboard end of the nacelle to an outboard end of the nacelle.

公开(公告)号：US20210062757A1

公开(公告)日：2021-03-04

申请号：US16985771

申请日：2020-08-05

Applicant: ROLLS-ROYCE plc

Inventor： Chia Hui LIM ,  Christopher T J SHEAF

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

Abstract: A gas turbine engine for an aircraft has an engine core, fan, air intake and gearbox. The engine core has a turbine, compressor, and core shaft connecting them. The fan is upstream of the engine core and has fan blades, the fan having 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 from 0.20 to 0.60. The gearbox receives an 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 gas turbine engine has a bypass ratio greater than 10; and the air intake defines highlight, throat and diffuser areas, wherein the gas turbine engine has a contraction ratio from 1.10 to 1.35, the contraction ratio being the ratio of the highlight area to the throat area.