Abstract: A gas turbine engine for an aircraft comprises: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a core casing surrounding the engine core. The gas turbine engine further comprises a core cowl surrounding the engine core and the core casing. The gas turbine engine further comprises an engine accessory gearbox driven by a take-off from the core shaft. The gas turbine engine further comprises an oil system having one or more oil pumps powered by the engine accessory gearbox for circulating lubricating oil around components of the engine including the engine accessory gearbox, and having an oil tank for receiving and storing oil scavenged from the engine components before recirculation thereto.

Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Abstract: A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core.

Abstract: An aircraft defining longitudinal, lateral and vertical directions the aircraft comprising: a main wing and a tail, each being pivotable about the lateral direction (B); a plurality of main propellers mounted to the main wing, and configured to pivot with the main wing; at least one cruise propeller mounted to the tail, and configured to pivot with the tail; each main propeller being stowable from a deployed position to a stowed position; wherein each main propeller has a fixed pitch, and each cruise propeller has a variable pitch.

Abstract: A method of operating a gas turbine engine compressor. The engine comprises a compressor having an environmental control system bleed port having an outlet in fluid communication with an aircraft environmental control system air duct, and an air turbine starter configured to rotate a compressor shaft of the gas turbine engine. The air turbine starter has an inlet in fluid communication with the environmental control system air duct via an air turbine valve. The method comprises determining a surge margin of the compressor, and where the surge margin of the compressor is determined to be below a predetermined minimum surge margin, opening the air turbine valve to supply air to the air turbine.

Abstract: There is disclosed a method of manufacturing a fan blade for a gas turbine engine, the method comprising: providing a root insert comprising quasi-isotropic short fibre reinforced resin, providing a first sub-laminate of a fibre-reinforcement pre-form for the fan blade on a first mould surface, placing the root insert on the first sub-laminate at a position corresponding to a root of the fan blade, providing a second sub-laminate of the pre-form over the root insert and the first-sub-laminate, so that the root insert is at an intermediate position between the first and second sub-laminates; and applying heat and pressure to form the pre-form.

Abstract: A thrust balancing mechanism for balancing axial loads on a rotor thrust bearing 3 is described. The mechanism comprises a piston arrangement 6 axially mounted on a stationary structure 2, about a center axis arranged, in use, in coaxial alignment with a rotating shaft 1 carrying the rotor thrust bearing 3. A hydrodynamic thrust bearing 8 is mounted, in use, between the piston 6 and the rotor thrust bearing 3. The piston 6 is pressurized so as to impart to the rotor thrust bearing 3, via the hydrodynamic thrust bearing 8, an axial load which counters an axial load imparted to the rotor thrust bearing 3 by the rotating shaft 1.

Abstract: A blade or a vane for a gas turbine engine. The blade or vane having an aerofoil and a leading edge member attached to the aerofoil. The leading edge has an electrically conductive support member and a nano-coating formed on the support member.

Abstract: An aircraft gas turbine engine nacelle comprises a thrust reversal arrangement. The thrust reversal arrangement comprises at least first and second circumferentially spaced fixed thrust reverser cascade boxes each comprising a plurality of thrust reverser vanes configured to direct air forwardly and circumferentially and at least one inter-leaved translating circumferential turning vane configured to direct air in a direction having a circumferential component. The circumferential turning vane is moveable from a stowed position provided between the first and second circumferentially spaced thrust reverser cascade boxes, and a deployed position axially rearwardly of the thrust reverser cascade boxes.

Abstract: A method of designing (300) an optimised component (100) for a load-bearing application comprises the steps of: analysing (302) a solid component (200) having an external shape to determine a stress distribution in the solid component (200) in response to an applied load; and defining (304), based upon the determined stress distribution in the solid component (200), a structure for an optimised component (100) comprising an outer skin layer (108) and an internal filament structure (110). The outer skin layer (108) forms an external shape substantially identical to the external shape of the solid component (200) and the internal filament structure (110) is enclosed by the outer skin layer (108), such that the optimised component (100) is configured to bear the applied load.

Abstract: A propulsion system includes a main gas turbine engine adapted for generating propulsive thrust during subsonic and supersonic flight operations and a supplementary propulsion unit adapted for generating additional thrust. The supplementary propulsion unit has an air intake and an exhaust for gas accelerated by the supplementary propulsion unit to provide the additional thrust and is adapted to generate the additional thrust during a limited range of subsonic flight operations, and to be dormant during other flight operations. The propulsion system has housing for the supplementary propulsion unit, including intake and exhaust covers which are moveable between deployed and stowed configurations. During the limited range of subsonic flight operations the intake and exhaust cover are moved to the deployed configuration to open the intake and the exhaust. During other flight operations the intake and exhaust cover are moved to the stowed configuration to close the intake and the exhaust.

Abstract: A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Abstract: A fan unit for a turbofan engine includes a plurality of fan blades mounted to a fan disc unit and a plurality of vortex-generating elements each of which is arranged to generate a respective vortex extending axially between a respective pair of adjacent fan blades during rotation of the fan unit about its rotation axis. During operation of a turbofan engine comprising the fan unit, fan-wake-induced mechanical degradation of compressor blades in a compressor downstream of the fan unit is eliminated, or the rate at which such degradation occurs is reduced.

Abstract: An aircraft propulsion system. The system comprises an internal combustion engine configured to drive a first propulsor and an electric generator. The system further comprises an electric motor or motors each configured to be powered by electric power provided by the generator, and configured to drive the first propulsor and/or a second propulsor. An electrical energy storage unit is provided, and is configured to store electrical energy supplied from the generator and supply electrical power to the electric motor. At maximum take-off power, the electric motor provides between 5% and 20% of total maximum aircraft propulsive take-off power.

Abstract: A low drag surface is provided for a fluid washed object, the low drag surface comprising an aerodynamic surface comprising a cut-out region, and a continuously translatable surface comprising a surface portion. The surface portion is positioned in the cut-out region such that the aerodynamic surface and the surface portion form a fluidwash surface, and the surface portion is translatable relative to the aerodynamic surface.

Abstract: The present disclosure relates to outlet guide vanes in a gas turbine engine, and in particular to such vanes with particular ranges of relative dimensions. Example embodiments include a gas turbine engine (10) comprising a plurality of outlet guide vanes (31) each having a length extending across a bypass duct (22) of the gas turbine engine (10), wherein for each outlet guide vane a minimum thickness to chord ratio is less than 80% of a maximum thickness to chord ratio.

Abstract: A vertical take-off and landing aircraft is shown. The VTOL aircraft has a main wing having a left wing and a right wing configured as folding wings, and one or more of a foreplane, having a left canard and a right canard configured as folding wings, and/or a tailplane, having a left stabiliser and a right stabiliser configured as folding wings. Each one of the folding wings has a fixed inboard section and a folding outboard section. The folding outboard section is downwardly foldable to a landing condition to support the aircraft on a surface.

Abstract: A gas turbine engine combustion chamber arrangement includes an annular outer wall and annular inner wall including an upstream and a downstream row of tiles. The downstream end of each tile in the upstream row of tiles has a rail extending towards and sealing with the outer wall and a lip extending in a downstream direction towards the tiles in the downstream row of tiles. The outer wall has a row of apertures to direct coolant onto the lips of the tiles in the upstream row of tiles. The downstream end of each tile in the upstream row of tiles has a plurality of fences extending in a downstream direction from the rail and each fence extends from the outer surface of the lip towards the outer wall. The row of apertures in the annular outer wall is arranged to direct coolant onto the lip between two circumferentially adjacent fences.

Abstract: Apparatus for a gas turbine engine, the apparatus comprising: a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing; a plurality of guide vanes extending radially from the outer wall of the core engine casing; a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; and an accessory gear box positioned within the second cavity of the torque box.

Abstract: This invention concerns an aircraft propulsion system in which an engine has an engine core comprising a compressor, a combustor and a turbine driven by a flow of combustion products of the combustor. At least one propulsive fan generates a mass flow of air to propel the aircraft. An electrical energy store is provided on board the aircraft. At least one electric motor is arranged to drive the propulsive fan and the engine core compressor. A controller controls the at least one electric motor to mitigate the creation of a contrail caused by the engine combustion products by altering the ratio of the mass flow of air by the propulsive fan to the flow of combustion products of the combustor. The at least one electric motor is controlled so as to selectively drive both the propulsive fan and engine core compressor.