Rolls-Royce Corporation

Rolls Royce Plc (Derby)

Richard Geoffrey Stretton

Michael Christopher Willmot

Kenneth Franklin Udall

Nicholas Howarth

David Michael Beaven

Alan Robert Maguire

Steven Aleksy Radomski

Peter Kevin Beardsley

Peter Beecroft

Angus Roy Smith

Nicholas Grech

Stephane Michel Marcel Baralon

Philip Geoffrey Woodrow

Tim O'Hanrahan

Ewan Fergus Thompson

Julian Mark Reed

Martyn Richards

Christopher A Murray

Adam MacGregor Bagnall

Christopher T J Sheaf

Alastair D Moore

John R Whurr

Richard Sharpe

Kenneth John Mackie

Richard V Brooks

Clare L Pool

Emlyn Jones

Pawel R Zakrzewski

Chia Hui Lim

John Richard Whurr

Clare Pool

Christopher Sheaf

Steven Radomski

Adam Macgregor Bagall

There is provided a gas turbine engine comprising a blower system for supplying pressurised air to an airframe via an airframe port. The blower system comprises a compressor configured to receive air from a bypass duct or a core of the gas turbine engine and to discharge compressed air into a delivery line extending from the compressor to the airframe port. The blower system also comprises a heat exchanger configured to transfer heat from the compressed air to a coolant and a valve arrangement configured to switch between operation of the blower system in a baseline mode and a cooling mode, the valve arrangement being configured to: selectively divert the compressed air within the delivery line to the heat exchanger for operation in the cooling mode; and/or selectively provide the coolant to the heat exchanger for operation in the cooling mode.

Gas turbine engine

A gas turbine engine includes an engine core including a compressor, a combustor, and a turbine, the compressor being connected to the turbines through a respective shaft; and a cabin blower system comprising: an electric variator comprising a first electrical machine connected to a first shaft arranged along a first axis, a second electrical machine connected to a second shaft arranged along a second axis, and a power management system; a cabin blower comprising a compressor driven by a third shaft arranged along a third axis, the compressor comprising an air inlet and an air outlet; and a differential gearbox. The gas turbine engine further includes an accessory gearbox arranged within an accessory gearbox casing and adapted to drive the cabin blower system.

Modular cabin blower system for aircraft

A gas turbine engine () for an aircraft comprises an engine core () comprising a turbine (), a compressor (), a core shaft (), and a core exhaust nozzle (), the core exhaust nozzle () having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (); a fan () comprising a plurality of fan blades; and a nacelle () surrounding the fan () and the engine core () and defining a bypass duct (), the bypass duct () comprising a bypass exhaust nozzle (), the bypass exhaust nozzle () having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

Turbine engine core and bypass flows having a defined fan-turbine radial distance

A gas turbine engine includes a support structure for attaching the engine to an aircraft pylon. The support structure includes: an engine-side interface member, a pylon-side interface member interfacing to the engine-side interface member, and a top V-shaped connection formation above the engine core and pair of side V-shaped connection formations on opposite lateral sides of the engine core, each V-shaped connection formation being formed by a pair of connection members meeting at a vertex, the vertex of the top V-shaped connection formation joining to the top of the engine-side interface member, the vertices of the side V-shaped connection formations respectively joining to the bottom ends of the engine-side interface member, and the connection members extending forwardly from their respective vertices to join to front fixation points at the core casing.

Support structure for attaching a gas turbine engine to an aircraft pylon

A gas turbine engine for an aircraft having an engine core configured with a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, the fan comprising a plurality of fan blades, with a nacelle surrounding the gas turbine engine, and a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle. An outer wall axis is defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle. An outer bypass duct wall angle is defined as the angle between the outer wall axis and the centreline, and the outer bypass duct wall angle is in a range from −15 to −2.5 degrees.

Fan arrangement for a gas turbine engine

A support structure for attaching an engine to an aircraft pylon at front, mid and rear attachment positions thereof, including a front mount joined to the engine and configured to attach to the pylon at the front attachment position and a rear mount joined to a core casing to attach to the pylon at the rear attachment position, each of the front and rear mounts configured to transfer lateral and vertical loads from the engine to the pylon, and the rear mount being spaced from the front mount such that yaw and pitch torques are transferred from the engine to the pylon through the front and rear mounts. The support structure also includes an axial load transfer formation to transfer axial loads from the engine to the pylon and a roll-torque transfer formation to transfer roll torque from the core casing to the pylon.

A seal assembly for a gas turbine engine having a rotor arranged to rotate about an axis in use. The seal assembly has a static support structure for the gas turbine engine and a casing structure of the engine. Rotation of the engine rotor causes a deflection of the casing structure relative to the static support structure in a first direction. A seal is provided at an interface between the static support structure and the casing structure, and comprising a first seal portion and a second seal portion spaced from one another in the first direction. The first seal portion is provided against a first surface of the casing structure and the second seal portion is provided against a second surface of the casing structure opposing the first surface. In an at-rest state in which the engine is not operational, the first and second surfaces are offset from an equilibrium position with respect to the static support structure such that there is a difference in compression of the first seal portion and the second seal portion between the static support structure and the casing structure. The offset is in a direction opposite to the first direction.

Seal assembly

A gas turbine engine comprising a planetary gear train, and a core engine casing. The gear train has a ratio of greater than approximately 3.0, with an input to the gear train being operatively connected to the compressor section, and an output from the gear train being operatively connected to the fan. The core engine casing encloses the compressor section and the turbine section. The fan has a diameter F, and the core engine casing has a diameter C. The core engine casing diameter C varies along an axial length of the core engine casing, and a ratio (C/F) of the core engine casing diameter C to the fan diameter F is within the range 0.2&#x3c;(C/F)&#x3c;0.4, along an axial length of the core engine casing.

Casing

A gas turbine engine () for an aircraft comprises a pylon attachment () and a shaft () defining an engine centreline (). The engine centreline lies in an engine central plane () which intersects the pylon attachment. The gas turbine engine comprises an intake () having a non-axisymmetric geometry and a medial plane () 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.

A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Turbine engine

A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

Large-scale bypass fan configuration for turbine engine core and bypass flows

A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

A gas turbine engine casing is described as having a cowl door hinged to a casing support structure by at least one hinge. The cowl door is openable outwardly from the casing to expose a casing interior. The hinge is located above a longitudinal axis of the casing and comprises a pivoting linkage arranged such that, upon actuation between closed and open cowl door conditions, the pivoting linkage moves an upper portion of the cowl door downwards towards the longitudinal axis.

Growing community of inventors

Loughborough, United Kingdom

Richard Geoffrey Stretton