Abstract: Apparatus for moving one or more processing devices relative to an object to be processed, comprising: first and second mounting units configured to be secured in a fixed position relative to the object to be processed and between which one or more tension members are mounted under tension; one or more carriages respectively connected to the one or more tension members, each of the carriages configured to accommodate a respective processing device; and a drive unit configured to move the one or more tension members so as to move the one or more carriages relative to the object.

Abstract: The present invention provides a method and apparatus for firing a ceramic component e.g. a ceramic core, within a firing receptacle containing firing powder. The method comprises using a shaping element to shape and/or define an upper surface of a first portion of firing powder at a predetermined height within the firing receptacle. The ceramic component is subsequently placed on the upper surface of the first portion of firing powder e.g. within a depression formed using the shaping element and covered with a second portion of firing powder. The shaping element may comprise a gridded/meshed plate.

Abstract: A variable diffuser comprises a passage, at least two vanes disposed within the passage, and at least two pennies. The passage is defined between opposing disk faces of a hub and a tip. Each of the vanes comprises a body having a leading edge and a trailing edge. The body extends between the hub face and the disk face. Each of the pennies is coupled to a respective vane body near an edge of the penny and an actuator. Rotation of at least one penny changes the orientation of the respective vane relative to the hub face.

Abstract: The novel nickel-base superalloy useful in an additive manufacturing process or a powder-based manufacturing process includes the following composition in wt %: Cr 8.0-8.5; Co 9.0-9.5; Mo 0.4-0.6; W 9.3-9.7; Ta 2.9-3.6; Al 4.9-5.6; Ti 0.2-1.0; Hf 0-0.05; C 0.005-0.03; B 0.005-0.02; Zr 0.005-0.1; Nb 0.2-1; Mn 0-0.6; and S 0-0.002 (?20 ppm); the balance nickel and incidental elements and unavoidable impurities.

Abstract: A bearing arrangement comprising a shaft with a recess in its end, a bearing housing, an annular first bush comprising an inner first portion and an inner second portion and a second bush that engages the recess of the shaft. The first bush bears against the bearing housing. The shaft engages the inner first portion of the first bush. The second bush engages the inner second portion of the first bush.

Abstract: A fuel injector provided for gas turbine engine has a nozzle including pilot fuel and mains fuel discharge orifice's for respectively spraying pilot and mains fuel flows into a combustor of the engine. The injector further has a feed arm extending to the nozzle for feeding fuel to the pilot and mains discharge orifices from pilot and mains supplies respectively. The nozzle and the feed arm contain one or more pilot passages for flow of the pilot fuel flow from the pilot supply to the pilot discharge orifice and one or more mains passages for flow of the mains fuel flow from the mains supply to the mains discharge orifice. The nozzle and the feed arm contain a plurality of the pilot passages and/or a plurality of the mains passages. The pilot and the mains passages intertwine repeatedly, and are both smoothly curved as they intertwine around each other.

Abstract: Turbomachinery comprising first and second sets of rotors configured to operate on a working fluid. The machinery further comprises first and second sets of electric machines coupled to the respective first and second rotors, and a coupling arrangement arranged to couple adjacent rotors of the first and second rotor sets to provide for fixed ratio, contra-rotation of the first and second rotor sets.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the turbine that drives the fan. The contribution of the turbine noise emanating from the rear of the engine to the Effective Perceived Noise Level (EPNL) is in the range of from 7 EPNdB and 30 EPNdB lower at a take-off lateral reference point than at an approach reference point.

Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the fan. The fan tip relative Mach Number at take-off at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is below 1.09. This results in low fan noise, along with optionally enabling a reduction in noise attenuation material.

Abstract: A propulsion system is disclosed comprising a gas turbine engine and an acceleration schedule which determines the rate of acceleration of the gas turbine engine from an idle condition in response to a demand for increased thrust off-idle. The acceleration schedule determines the rate of acceleration in dependence upon the value of an engine parameter of the engine the value of which is substantially unaltered by variation in the magnitude of an electrical load drawn from the engine while it is operating in the idle condition.

Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), and an outer core casing provided outwardly of compressor blades, the inner core casing and outer core casing defining a core working gas flow path (B) therebetween; a fan (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) spaced radially outwardly from the first outer core casing, and wherein at an axial plane (E) of an inlet to the high pressure compressor, the second outer core casing has an inner radius at least 1.4 times the inner radius of the first outer core casing.

Abstract: A gas turbine engine comprises a gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a carrier. The sun gear meshes with the planet gears and the planet gears mesh with the annulus gear. A planet gear comprises a first annular structure, a plurality of teeth extending radially from the first annular structure and a second annular structure abutting and secured to a cylindrical surface of the first annular structure. The first annular structure and the teeth consist of steel, the second annular structure comprises a metal matrix composite, the metal matrix comprising aluminium, aluminium alloy, steel, titanium or titanium alloy, the reinforcing material comprising titanium diboride, titanium carbide or titanium nitride. The second annular structure reduces the mass but maintains, or increases, the stiffness of the planet gear to carry centrifugal loads and to maintain roundness of the planet gears.

Abstract: There is disclosed a blade for a gas turbine engine comprising a blade root that is elongate in a longitudinal direction. The blade root comprises a first flank surface and a second flank surface on opposite sides of the blade root that are asymmetrical about a longitudinal mid-plane of the blade root.

Abstract: A blade for a gas turbine engine comprises an aerofoil body having a suction side, a pressure side, and a trailing edge. An internal cooling passageway is provided in the aerofoil body, and an ejection slot in fluid communication with the cooling passage and provided at the trailing edge of the aerofoil body. The ejection slot is defined between a pressure side wall and a suction side wall. Both the suction side wall and the pressure side wall include a mid-section and a trailing edge section adjacent the mid-section, and the thickness of the suction side wall and the pressure side wall reduces to define a taper with a wedge angle less than or equal to 20 degrees.

Abstract: An aircraft has an internal combustion engine and a wing including a hollow structural member. The aircraft has an electrical network including: an electrical motor configured to drive an aircraft propulsor; at least one conductor configured to electrically couple the electrical motor and an electrical storage device; wherein the electrical conductor is formed of the hollow structural member of the aircraft wing.

Abstract: A highly efficient gas turbine engine includes a fan which is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

Abstract: A method (56) for fabricating an integral assembly (24) is disclosed. The method comprises providing (56) a first workpiece (26) having a first surface (46) and a second workpiece (28) having a second surface (50), performing a first bonding process between the first surface (46) of the first workpiece (26) and the second surface (50) of the second workpiece (28) to form a sub-assembly workpiece (48), and performing a second bonding process between a third surface (52) of the sub-assembly workpiece (48) and a fourth surface (54) of a third workpiece (30) to form the integral assembly (24). The material properties of the first workpiece (26) are different from material properties of the third workpiece (30).

Abstract: A method is provided of controlling a gas turbine having a shaft connecting a compressor to a turbine, as well as having a reheat system, and a gas turbine. The method includes the steps of: operating the engine using the reheat system to provide a mass flow rate of reheat fuel into a gas flow of the gas turbine engine downstream of an exit of the turbine; detecting a shaft break event in the shaft; and in response to this detection, maintaining the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit, whereby the maintained mass flow rate of reheat fuel raises a back pressure downstream of the turbine and thereby reduces a rotational speed of the turbine.

Abstract: A combustor for a gas turbine engine comprising: a combustion chamber; at least one fuel spray nozzle operable to deliver a fuel-air mixture into the combustion chamber, wherein during operation of the gas turbine engine the fuel-air mixture is combusted in the combustion chamber, thereby producing a combustion flame; a microwave generator coupled to a waveguide arranged to guide microwaves from the microwave generator into the combustion chamber such that the microwaves are incident on at least a portion of the combustion flame; and a detector operable to detect at least a portion of the microwaves reflected by the combustion flame and/or atomised fuel droplets.

Abstract: An insert fixture has a base, a plurality of mounting brackets, and a plurality of separators. The plurality of separators extends vertically from the base and includes a plurality of grid portions extending the length of the insert fixture and a plurality of divider portions, which connect to the plurality of grid portions to form a plurality of individual component holders around one of the plurality of mounting brackets. Each individual component holder has two separated grid portion sections positioned on either side of the bracket. These grid portions have two divider portions which are also separated and positioned either side of the bracket at an angle relative to the two grid portions. The individual component holder forms a cell around the mounting bracket. The insert fixture may be constructed from a molybdenum alloy, lanthanum oxide and/or titanium zirconium molybdenum.