Abstract: A thermal management system comprises a first heat exchanger configured to exchange heat between a first component and a first working fluid, a first working fluid compressor downstream in first working fluid flow of the first heat exchanger and configured to compress the first working fluid, a second heat exchanger downstream in first working fluid flow of the compressor and configured to exchange heat between the first working fluid and a second working fluid and an expander downstream in first working fluid flow of the second heat exchanger, and configured to expand and cool first working fluid and deliver cooled first working fluid to the first heat exchanger. The system further comprises a third heat exchanger upstream in second working fluid flow of the second heat exchanger, and configured to exchange heat between a second component and the second working fluid.

Abstract: An aircraft cabin blower system includes: a transmission configured to receive mechanical power from a first part of a gas turbine engine in the form of a first transmission input; and an electrical circuit including a first electrical machine, a second electrical machine, and a power management system, wherein an output of the transmission is configured to drive a cabin blower compressor when operating in a blower mode, the first electrical machine being configured to receive mechanical power from a second part of the gas turbine engine and act as a generator to provide electrical power to the power management system, and the second electrical machine being configured to act as a motor providing mechanical power to the transmission in the form of a second transmission input, the second electrical machine being driven by electrical power from the power management system.

Abstract: A method of heat treating a localised region of a component including providing a heat treatment assembly with two or more subassemblies. Each subassembly is configured to partially circumscribe a portion of the component for heat treatment. Each subassembly includes a housing configured to partially circumscribe a portion of the component. An insulator and a heater are provided within the housing and extend to partially circumscribe a component and are arranged such that the insulator is between a wall of the housing and the heater. The method includes positioning the subassemblies adjacently around the component so that the subassemblies fully circumscribe the component. The subassemblies are connected together. The method includes activating the heater to heat the component in a region adjacent the heater.

Abstract: A boundary layer propulsor comprises a rotor and a plurality of first aerofoil blades. The rotor has an axis of rotation. The plurality of first aerofoil blades extends radially from the rotor and is arranged in a circumferential array around the axis of rotation. Each of the first aerofoil blades has, in a radially outward sequence, a radially proximal portion, a middle portion, and a radially distal portion. The radially proximal portion has a first cambered cross-section, the middle portion has a second uncambered cross-section, and the radially distal portion has a third cambered cross-section. The first cambered cross-section is cambered in an opposite sense to the third cambered cross-section.

Abstract: A computer implemented method of determining a dimension of a gap between an edge of an aerofoil and a surface of an engine casing, the method comprising: receiving data; generating a three dimensional model of the surface of the engine casing from the received data; identifying the edge of the aerofoil in the received data; determining a three dimensional position of a first location along the edge of the aerofoil in the received data using the identified edge; and determining a distance between the determined three dimensional position of the first location and the three dimensional model of the surface of the engine casing using an algorithm.

Abstract: A bearing support structure for a gas turbine engine located within an internal portion of the engine. The bearing support structure has a plurality of stators, a first section, a second section, a first bearing assembly, and a second bearing assembly. The first section depends forwardly from the plurality of stators relative to the longitudinal axis. The section second depends rearwardly from the plurality of stators relative to the longitudinal axis and is detachably mounted to the plurality of stators. The first bearing assembly is supported relative to the plurality of stators by the first section. The second bearing assembly is supported relative to the plurality of stators by the second section. The second section is detachably mounted to the plurality of stators.

Abstract: A core duct assembly for a gas turbine engine, the core duct assembly including: a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.

Abstract: A method of removing a ceramic coating from a ceramic coated metallic article without damaging the metallic bond coating, the metallic article having a first and second portions, each of the portions comprising a metallic bond coating and a ceramic coating on the metallic bond coating, the ceramic coating on the second portion being less porous than the ceramic coating on the first portion. The method comprises the steps of a) immersing the ceramic coated metallic article in a caustic solution; b) maintaining the ceramic coated metallic article in the caustic solution at atmospheric pressure for a predetermined time period and at a predetermined temperature; c) removing the ceramic coated metallic article from the caustic solution; d) rinsing the ceramic coated metallic article in water at ambient temperature; e) water jet blasting the first portion of the metallic article to remove the ceramic coating; and f) water jet blasting the second portion of the metallic article to remove the ceramic coating.

Abstract: Apparatus for selective delivery of pressurised gas comprises a containment wall which at least partially bounds of a volume of pressurised gas during use of the apparatus, and a valve directly coupled to an output port of the containment wall such that the containment wall and the valve retain the pressurised gas when the valve is closed. The apparatus obviates the need for a duct coupling the output port to the input of the valve, and hence reduces the number of interfaces between the output port and the input of the valve from two, as in arrangements of the prior art, to one. The apparatus presents fewer failures modes compared to the case where a duct couples the output port to the valve.

Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting them. The engine includes a fan, with a plurality of fan blades, located upstream of the core and a gearbox receiving an input from the core shaft and outputting drive so the fan is at a lower rotational speed than the core shaft. The turbine includes a plurality of stages of axially spaced rotor blades mounted on a rotor, which are surrounded by a turbine casing. The turbine has an inlet defined at an upstream end of a first stage of blades and an outlet defined at a downstream end of a last stage of blades and a ratio of the area of the outlet to the inlet is at between 2.5 and 3.5. This increases the pressure ratio of and power extracted from the turbine and the engine.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A single-piece gas turbine engine bleed duct for a gas turbine engine including: a main airflow conduit configured to transmit a bleed flow to a location outside the gas turbine engine; a pressure regulating valve for regulating airflow through the main airflow conduit; a first inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the first inlet duct including a non-return valve; and a second inlet duct for directing airflow to the main airflow conduit and toward the pressure regulating valve, the second inlet duct including a control valve; wherein each of the first and second inlet ducts are directly connectable to an engine casing of a gas turbine engine.

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine comprising: a compressor; a first turbine; and a first compressor bleed valve in fluid communication with the compressor and configured to release bleed air from the compressor; wherein the first compressor bleed valve is configured to release bleed air to a downstream location in the engine, the downstream location being downstream of the first turbine; wherein the first compressor bleed valve is configured to open wherein the first compressor bleed valve is configured to open to at least two positions, to thereby release a variable amount of bleed air from the compressor.

Abstract: A gas turbine engine comprising an engine core comprising a compressor; a compressor bleed valve in fluid communication with the compressor and configured to release bleed air from the compressor; and a combustor comprising a fuel manifold configured to provide fuel to the combustor; wherein the fuel manifold is in thermal contact with a cooling conduit; and the gas turbine engine further comprises a fluid conduit to supply bleed air from the compressor bleed valve to the cooling conduit.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A method of forming a diffusion bonded joint comprises the steps of: providing a first component having a first faying surface; providing a second component having a second faying surface; applying a lamellar coating to at least one of the first faying surface and the second faying surface; and bringing the first and second faying surfaces into contact in a diffusion bonding operation to form the diffusion bonded joint.

Abstract: A method of separating a gas turbine engine includes steps to support the gas turbine engine on a core stand, and to move a fan stand into axial alignment with a core stand. The fan stand includes a base frame and a fan case frame coupled to the base frame at one edge by a hinge. The fan case frame rotates about an axis of the hinge between abutting the base frame and being perpendicular to the base frame, and includes a coupling arrangement that couples a fan case to the fan case frame. The method further includes steps to rotate and tilt the fan case frame into abutting relation with the fan case, couple the fan case to the fan case frame, decouple the fan case and core engine, and translate at least part of the core stand axially to separate the gas turbine engine.

Abstract: A method of manufacturing a component, the method including a maraging steel blank with an initial shape; and performing an incremental cold forming operation on the maraging steel blank, wherein the incremental cold forming operation reduces a thickness of the maraging steel blank.

Abstract: A shaft apparatus for a gas turbine engine comprising: a first shaft portion; a second shaft portion; and a ratchet mechanism configured to permit the first shaft portion to rotate with respect to the second shaft portion in a first direction, and to prevent the first shaft portion from rotating with respect to the second shaft portion in a second direction opposite to the first direction. A gas turbine engines comprising the shaft apparatus and methods of operating a gas turbine engine are also disclosed.