Abstract: A method of manufacturing an induction motor rotor assembly, the method includes the steps of: providing a rotor; machining a plurality of re-entrant slots axially along an outer surface of the rotor; positioning a sleeve concentrically over the outer surface of the rotor; applying a friction stir welding process to the sleeve along each re-entrant slot axially along the outer surface of the rotor to cause the sleeve material to plasticise and flow into the axial re-entrant slot to form an axial re-entrant slot bar; and providing an electrical connection at each of the opposing axial ends of the rotor between respective ones of opposing ends of each of the axial re-entrant slot bars to thereby form the induction motor rotor.

Abstract: A gas turbine engine has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The bypass flow rate at the optimum bypass efficiency is appreciably lower than the maximum bypass flow rate at the given conditions. This results in increased design flexibility and improved overall engine performance.

Abstract: A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine system comprising one or more turbines (17, 19), a compressor system comprising one or more compressors (14,15), and a core shaft (26) connecting the turbine system to the compressor system, wherein a compressor exit pressure (P30) is defined as an average pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core (11) further comprises an annular splitter (70) at which flow is divided between a core flow (A) that flows through the engine core and a bypass flow (B) that flows along a bypass duct (22), wherein stagnation streamlines (110) around the circumference of the engine (10), stagnating on a leading edge of the annular splitter (70), form a streamsurface (110) forming a radially outer boundary of a streamtube that contains all of the core flow (A); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan bl

Abstract: A gas turbine engine includes: an engine core, compressor system, and core shaft. A compressor exit pressure is defined as an average airflow pressure at the exit of the highest pressure compressor at cruise conditions. The core has an annular splitter and bypass flow. Stagnation streamlines around the engine circumference form a streamsurface. A fan is upstream the core with blades having leading and trailing edges, and a radially inner portion within the streamtube. A fan root entry pressure is an average airflow pressure across the radially inner portion leading edge of each fan blade at cruise conditions. An overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure. A bypass jet velocity is defined as the jet velocity of air flow exiting the bypass exhaust nozzle at cruise conditions. A jet velocity is in a range between 4.7 m/s and 7.7 m/s.

Abstract: Apparatus for controlling a power generation system, the apparatus comprising a controller configured to: identify a trigger indicative of a future change in electrical power output by the power generation system to a first power level; control the power generation system to change electrical power output to a second power level in response to the trigger, the second power level being equal to, or different to the first power level; and control supply of at least a portion of the electrical power output from the power generation system at the second power level to an electrical energy storage system to charge the electrical energy storage system.

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is upstream of the core and includes a plurality of fan blades, and has 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 0.20 to 0.60 and defines highlight, throat and diffuser areas. The nacelle at least partially surrounds the core and fan. The gearbox receives 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 engine has a bypass ratio greater than 10. The nacelle has a length and the ratio of the length of the nacelle to the fan diameter is 0.4 to 2.5.

Abstract: A method of controlling a gas turbine engine including receiving an instantaneous thrust demand for current operation of the engine, determining the inlet flow rate and/or the pressure ratio within the compressor of the engine and determining whether the inlet flow rate and/or the pressure ratio match the working line for the compressor. The angle of one or more vane of the compressor is adjusted according to a closed control loop if the inlet flow rate and/or pressure ratio lie outside said desired range in order to adjust the inlet inflow rate and/or pressure ratio to meet the working line. The fuel flow to the engine combustor is adjusted concurrently in order to meet the thrust demand.

Abstract: A method and system for performing computational jobs securely on a shared computing resource. Data files for the computational job are encrypted on a secure system and the encrypted data files are stored in a data store on the shared computing resource. A key distribution server is established using a secure enclave on a front end of the shared computing resource. Cryptographic keys and application binaries are transferred to the enclave of the shared computing resource using a session key. The computational job is run using an application launcher on compute nodes of an untrusted execution environment of the shared computing resource, the application launcher obtaining the application binaries and the cryptographic keys from the key distribution server.

Abstract: A gas turbine engine for an aircraft includes an engine core, a fan, an air intake and a gearbox. The engine core includes a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan is upstream of the engine core and includes a plurality of fan blades, and has 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 0.20 to 0.60. The gearbox receives an input from the core shaft and drives 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 a highlight area, wherein the ratio of highlight area to throat area is from 1.15 to 1.35 and the ratio of diffuser area to throat area is from 0.85 to 1.15.

Abstract: Cladding of the interior of a component part of a pressure vessel is shown. A lining which conforms to at least a portion of the interior geometry of the component is positioned in the interior of the component. The lining is then pressed into the component past its yield strength. The lining is then fused to the component.

Abstract: A gas turbine engine for an aircraft includes an engine core, a fan, an air intake and a gearbox. The engine core includes a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan is upstream of the engine core and includes a plurality of fan blades, the fan having a diameter greater than 2.0 m. The air intake is located upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60. The gearbox receives an input from the core shaft and outputs drive 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 a highlight area, a throat area and a diffuser area, wherein the ratio of the throat area to fan face area is from 0.94 to 1.05.

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake and gearbox. The engine core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is located upstream of the core and includes a plurality of fan blades, the fan having a diameter greater than 2.0 m. The air intake is located upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The gearbox receives an input from the core shaft and outputs drive to the fan so as 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 has an intake length and the ratio of the intake length to the fan diameter is from 0.20 to 0.60.

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.

Abstract: An oil delivery system is provided that includes a tank for a gas turbine engine, where the tank is positioned radially outward from a compressor section, a combustor section, and/or a turbine section of the gas turbine engine. The tank is configured to store oil for the gas turbine engine. The oil delivery system further includes a primary lubrication system including a sump of a power gearbox, a pump, and an oil feed line. The oil feed line extends from the tank to the primary lubrication system. The oil feed line is configured to allow a flow of oil to pass from the tank to the pump and from the pump through the power gearbox to the sump of the power gearbox.

Abstract: A gas turbine engine for an aircraft includes an engine core, fan, air intake, nacelle, and gearbox. The core includes a turbine and compressor, connected by a core shaft. The fan is upstream of the core and includes a plurality of fan blades, and has 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 0.20 to 0.60. The nacelle at least partially surrounds the core and fan. The gearbox receives 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 engine has a bypass ratio greater than 10. The engine has a local diffuser angle from 0 to 18 degrees, a local peak diffuser angle from 0 to 22 degrees, and/or a bulk diffuser angle from 0 to 15 degrees.

Abstract: An electrical machine has a stator carrying electrical windings which protrude at opposite ends of the stator to form respective, ring-shaped end windings. The electrical machine further has a rotor having a plurality of magnetic field-producing elements for producing a rotor magnetic field which interacts with a stator magnetic field produced by the windings. The electrical machine further has a coolant bath for holding liquid coolant. The electrical machine further has a heat sink in thermal contact with at least one of the end windings, the bath and the heat sink being configured such that the heat sink is immersed in the coolant held in the bath. The heat sink defines one or more fluid pathways configured such that vapour bubbles, formed when coolant in contact with the heat sink boils, escape by rising through the heat sink.

Abstract: A rotary abrasive machining tool (101) is shown. The tool comprises a hub (103) having plurality of axially-oriented radial slots in the outer circumference thereof, and plurality of abrasive segments each of which is located in a respective slot in the hub and which form an abrading surface (102). The abrasive segments comprise a tab for location in a slot in the hub, and an abrading edge defining a plurality of abrasive elements.

Abstract: A gas turbine engine component made of a nickel-based superalloy, the gas turbine engine component comprising a protective coating. The protective coating includes an inner diffusion barrier layer including any one or any combination of elements selected from the group consisting of platinum, palladium, tantalum, tungsten, hafnium and iridium, and an outer layer of hard material formed of hard particles embedded in a matrix.

Abstract: A gas turbine is provided for an aircraft comprising an engine core and a core flow path, a fan, a front drum cavity arranged radially inward of the core flow path, and a front bearing chamber. The front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor. The front drum inlet is through a seal, and the front drum outlet is through a spaced gap.

Abstract: An additive layer manufacturing apparatus comprising a doctor blade (57) having a tip (51) which, in use, is arranged proximal to a material layer (59) to be recoated during an additive layer manufacturing method, the doctor blade (57) having a cavity (50) with an inlet and an outlet, the outlet located at the tip (51) and the inlet connectable to an air supply (53) whereby air can be directed through the cavity (50) to the outlet and onto a surface (59a) of the material layer (59).