Abstract: A gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a total electrical power PEM1 (W), and the gas turbine engine is configured to generate a total shaft power PSHAFT (W); and a ratio R of: R = ( Total ? Shaft ? Power = P SHAFT ) ( Total ? Electrical ? Power ? Generated = P EM ? 1 ) is in a range of between 0.005 and 0.020.

Abstract: An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.

Abstract: A gas turbine engine includes a core engine having a compressor, a combustor and a turbine in sequential air flow series. The engine further includes a fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit and a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit. At least first and second compressor bleed offtakes are at different pressure stages of the compressor, each being configured to bleed a portion of air from the compressor. At least first and second compressor bleed offtake valves are configured to control flow through the first and second bleed offtakes respectively. The burner is configured to receive bleed air from the compressor bleed offtakes.

Abstract: A gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Abstract: A gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Abstract: There is disclosed a method of manufacturing a composite component. A preform for the component is laid-up on a lay-up tool so that a first surface of the preform conforms to a lay-up profile of the tool. The preform is transferred to a forming tool comprising opposing first and second forming surfaces, each having a near net shape profile corresponding to a respective side of the component. The preform is formed in the forming tool to a near net shape of the component. The preform has a preform bulk in excess of the near net shape of the component prior to forming, and the lay-up profile is offset relative to the near net shape profile of the first forming surface to accommodate a proportion of the preform bulk. Accordingly, both the first surface and an opposing second surface of the preform displace during forming to conform to the respective near net shape profiles of the forming surfaces.

Abstract: A combined cycle engine is formed of a gas turbine cycle coupled with a fuel expansion cycle in which fuel for the gas turbine cycle is the working fluid of the fuel expansion cycle. The fuel expansion cycle comprises, in flow series, a fuel pump for pumping a cryogenic fuel for the gas turbine cycle, a first heat exchanger for heating the fuel, an expander to recover power by expansion of the fuel and to thereby drive a load, and a second heat exchanger to reheat the fuel. The gas turbine cycle comprises, in flow series, a compressor, a combustor for combustion of reheated fuel delivered from the second heat exchanger, and a turbine configured to drive the compressor. The first heat exchanger and the second heat exchanger are stationed downstream of the combustor and configured to transfer heat from combustion products to the fuel in the fuel expansion cycle.

Abstract: There is provided a dynamic sealing assembly for a rotary machine, comprising a primary sandwich plate, a secondary sandwich plate and a bristle pack. The primary sandwich plate comprises a plurality of primary vane openings, and the secondary sandwich plate comprises a plurality of secondary vane openings. The bristle pack comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings overlies and aligns with a respective secondary vane opening to form a vane channel for receiving a vane along a longitudinal axis of the dynamic sealing assembly. The bristle pack is configured to: provide a brush seal between each vane received within the respective vane channels and the dynamic sealing assembly; and allow relative movement between the dynamic sealing assembly and the vane received within each vane channel along the longitudinal axis.

Abstract: A method of fabricating a component is described and involves retaining a canister as an integral part of the component after performing an isostatic pressing process. The method comprises: providing a canister having a canister wall that encloses an internal cavity, the canister wall comprising at least a first wall section and a second wall section, where the first wall section and the second wall section are of different materials; filling the internal cavity with a powdered material; performing an isostatic pressing process on the filled canister to consolidate the powder; and retaining the canister as an integral part of the component such that an internal structure of the component comprises the consolidated powder and the canister wall forms at least part of a surface of the component that covers the internal structure.

Abstract: A method of operating an aircraft that includes a propulsion system. The propulsion system includes a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and controlling the propulsion system based on the at least one fuel characteristic.

Abstract: A method for scanning of an object in a scanning apparatus includes disposing the object and a complementary object on a support of the apparatus, so that the objects are positioned between an imaging beam emitting element and an imaging beam receiving element oppositely disposed to either side of the support. The support is rotatable relative to the emitting and receiving elements about an axis of rotation to allow creation of an image from projections each taken at a different relative angle. A volume of the complementary object is solid or filled with a filling material and the complementary object is configured to reduce the variation in imaging beam attenuation across the objects or the part to be scanned of the object at the multiple relative angles of rotation. The method includes operating the scanning apparatus at the multiple relative angles of rotation to produce an image of the object.

Abstract: A non-contact excitation system for a spin rig includes: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements include magnets; each exciter element being moveable between an active position, in which a magnetic force is provided between each exciter element and at least one of the test piece elements, and an inactive position, in which the magnetic force between the exciter magnets and the at least one test piece element is reduced.

Abstract: There is provided an air pressurisation system for an aircraft. The air pressurisation system comprises a blower compressor, a delivery line and a catalyst material. The blower compressor is configured to be mechanically coupled to a spool of a gas turbine engine and configured to receive an inlet flow of gases from a bypass duct of the gas turbine engine. The delivery line is configured to convey gases received from the blower compressor to an airframe port for supply to an airframe system. The catalyst material is disposed along the delivery line and configured to catalyse a reaction of volatile organic compounds within gases conveyed by the delivery line.

Abstract: An engine for an aircraft includes an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.

Abstract: A battery cooling block, including: a housing defining a flow chamber and including a first and second port configured to allow flow of coolant through flow chamber along a first axis; the housing further including a plurality of compartments disposed in the flow chamber, each compartment extending along a respective compartment axis perpendicular to first axis and each compartment configured to receive respective cylindrical battery cell; wherein each compartment is defined by partially open compartment wall within flow chamber, compartment wall including a plurality of wall segments extending parallel to compartment axis and which are circumferentially spaced apart around compartment axis wherein respective battery cell is partially exposed to flow of coolant. A battery assembly, including: plurality of cylindrical battery cells; and battery cooling block as described above.

Abstract: Described herein is a non-contact excitation system for a spin rig comprising: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements comprise magnets; and wherein, in use, the exciter elements are arranged around the test piece; wherein each exciter element is axially and radially offset from the plurality of test piece elements. Also described is a method of testing a test piece using the system.

Abstract: A ceramic matrix composite (CMC) component is provided that includes: a CMC body in which an environmental protection layer is completely embedded within a CMC material of the CMC body, the environmental protection layer comprising a ceramic that has a higher impact and/or environmental resistance than the CMC material. Methods for manufacturing the CMC component are also provided.

Abstract: A fuel management system for a gas turbine engine, the fuel management system comprising a fuel supply line configured to supply fuel from a system inlet to a combustor of the gas turbine engine via a combustor pump and combustor valve. A heat exchanger is configured to reject heat from a thermal load of the gas turbine engine to fuel in the fuel supply line between the system inlet and the combustor valve. There is a downstream recirculation line configured to recirculate a downstream excess portion of fuel from the fuel supply line, the downstream recirculation line extending from a downstream recirculation point on the fuel supply line between the heat exchanger and the combustor valve.

Abstract: A combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

Abstract: A fuel characteristic determination system and method for determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft. The system includes a sensor component that is formed from a nitrile seal material and includes a surface that can be exposed to the fuel; a sensor for measuring a swell parameter of the seal material; and a fuel characteristics determination module for determining one or more fuel characteristics of the fuel based on the swell parameter. The method includes: exposing one or more seals of a fuel system of an aircraft to fuel within the fuel system; exposing a sensor component, made from the same material as the one or more seals, to the fuel; and measuring a swell parameter of the seal material.