Abstract: An oil tank (100) for a gas turbine engine is provided. The oil tank (100) comprises an oil inlet (102), an oil outlet (104), and a body including a coiled portion (106) interposed between the oil inlet (102) and the oil outlet (104). Oil is received by the coiled portion (106). The coiled portion (106) acts to at least partially de-aerate oil received from the oil inlet (104).

Abstract: Combustion equipment includes annular combustion chamber having an annular upstream wall structure. Annular outer casing surrounds and is spaced from annular combustion chamber and annular inner casing is within and spaced from annular combustion chamber. Ogee shaped annular cowl is positioned upstream of annular upstream wall structure and a stage of compressor outlet guide vanes is positioned upstream of annular cowl. Pre-diffuser extends in downstream direction from stage of compressor outlet guide vanes, pre-diffuser is positioned upstream of cowl. Stage of compressor outlet guide vanes is connected to outer and inner casing. Outer fairing extends radially inwardly from outer casing such that annular outer casing, outer fairing and pre-diffuser define annular outer cavity and outer fairing is spaced from cowl or inner fairing extends radially outwardly from inner casing such that annular inner casing, inner fairing and pre-diffuser define an annular inner cavity and inner fairing is spaced from cowl.

Abstract: The disclosure relates to a gas turbine engine having an accessory gearbox assembly having an accessory gearbox. The accessory gearbox assembly is mounted adjacent a core of the gas turbine engine and has a support formation for supporting a conduit. The conduit extends between a first location on the engine and a second location on the engine. The first location and the second location may be spaced from the accessory gearbox such that the conduit does not operatively communicate with the accessory gearbox in use. The accessory gearbox may be axially mounted and may provide a bridge for supporting a plurality of conduits.

Abstract: An apparatus measuring thrust of an aircraft gas turbine engine includes a core shaft connecting a turbine and compressor, a fan and gearbox with a sun gear driven by the core shaft, a plurality of planet gears, an annulus gear mounted in a static structure, and a planet carrier driven by the fan via fan shaft. The apparatus includes a sensor to measure force applied by the annulus gear on the static structure and first and second sensors to measure rotational speed of the core and fan shafts. A processor determines restoring torque on the annulus gear from measurement of force applied by the gear on the static structure, torque applied to the fan by the planet carrier using rotational speeds of core and fan shafts and restoring torque on the annulus gear, and thrust of the fan from torque applied to the fan and the fan's rotational speed.

Abstract: A method is provided for fabricating an insulated metal coil for an electrical machine. The method includes 3D printing a metal coil having a plurality of turns. The method further includes subsequently infiltrating insulating material between the turns of the metal coil to electrically insulate the turns from each other.

Abstract: A brush seal for sealing a leakage gap in an axial flow path between a relatively higher fluid pressure region and a relatively lower fluid pressure region, that includes an outer housing and an inner housing located at least partially within and configured for radial displacement relative to the outer housing, wherein the upstream facing inner surface of the outer housing is free of any protrusion toward the first downstream facing outer contact surface and second downstream facing outer contact surface.

Abstract: A method of machining a surface of a component. The method comprises scanning the surface of the component to obtain scanned electronic 3D data representing the scanned surface of the component locating a surface defect of the scanned surface and identifying a defect region that surrounds and includes the surface defect, and providing electronic 3D data representing a patch having the desired shape of the defect region. The method also comprises transforming the patch to generate a tooling path for repairing the surface defect. The transformation comprising translating a plurality of nodes of the patch. The translation distance of each node based on the distance of that node from an origin node of the patch. The method further comprises machining the surface of the component according to the generated tooling path.

Abstract: A gas turbine engine (10) comprising: a high pressure turbine (17); a low pressure turbine (19); a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27); a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four or five compressor stages (14); the high pressure compressor (15) consists of eight or nine compressor stages; the low pressure turbine (19) comprises four or more stages; and the high pressure compressor (15) and low pressure compressor (14) together define a core overall pressure ratio of greater than 36:1.

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 the noise emanating from the front of the fan. The contribution of the fan noise emanating from the front of the engine to the Effective Perceived Noise Level (EPNL) 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 in the range of from 0 EPNdB and 12 EPNdB lower than the contribution of the fan noise emanating from the rear of the engine to the EPNL at the take-off lateral reference point.

Abstract: Fuel spray nozzle for generating a spray of atomised liquid fuel in a combustor of a gas turbine engine. The nozzle includes a flow circuit that has in flow series: a gallery that receives fuel flow, plural circumferentially-spaced restrictor passages arranged in a row around the nozzle, plural conditioning passages configured to impart a circumferential component to their respective portions of the fuel flow, and an annular spin chamber which forms a swirling fuel flow which is discharged at an exit port. The restrictor passages form flow restrictions which in use produce a pressure differential between the gallery and the spin chamber to evenly circumferentially distribute the fuel flow between the restrictor passages. The conditioning passages have increased flow cross-sectional areas relative to the flow cross-sectional areas of the restrictor passages, such that the restrictor passages produce substantially all of the pressure differential between the gallery and the spin chamber.

Abstract: An electromechanical system comprising: a rotor comprising a plurality of teeth spaced about a circumference of the rotor, wherein the teeth of the rotor are equally spaced about the rotor according to a tooth spacing angle, and are skewed in an axial direction such that the circumferential positions of the teeth of the rotor vary along the axial length of the rotor, and first and second stator segments, each extending partway about the rotor and comprising a field winding and poles arranged to magnetically interact with the teeth of the rotor such that an alternating current (AC) back-emf is induced in the field winding upon rotation of the rotor. The poles of the first stator segment are angularly displaced about the rotor from the poles of the second stator segment such that the back-emf induced in the field winding of the first stator segment is phase shifted with respect to the back-emf induced in the field winding of the second stator segment.

Abstract: A gas turbine engine comprising: an inner core nacelle; an outer fan nacelle; a bypass duct between the inner core nacelle and the outer fan nacelle; at least one bifurcation that extends between the inner core nacelle and the outer fan nacelle; and a cooling system, wherein the cooling system comprises at least one pipe for conveying a fluid to be cooled, the at least one pipe forming part of a fluid system of the engine, wherein the at least one pipe passes through the at least one bifurcation, and wherein at least a portion of one or more of the pipes is arranged to bring the fluid to be cooled into a heat exchanging relationship with a fluid, e.g. air, flowing in the bypass duct.

Abstract: A gas turbine engine for an aircraft including an engine core including a turbine, compressor, and core shaft connecting the turbine and compressor; a fan located upstream of the engine core including a plurality of fan blades; a gearbox that can receive input from the core shaft and can output drive to the fan at a lower rotational speed than the core shaft, an epicyclic gearbox including a sun gear, planet gears, a ring gear, and a planet carrier on which the planet gears are mounted; and a gearbox support. A first gearbox support shear stress ratio: torsional ? ? shear ? ? stress ? ? of ? ? the ? ? gearbox ? ? support ? at ? ? maxiumum ? ? take ? ? off ? ? conditions radial ? ? bending ? ? stiffness ? ? of ? ? the ? ? gearbox ? ? support is less than or equal to 4.

Abstract: An assembly of a servo pump and a hydraulic motor. The assembly has a housing which contains the pump and the motor. The motor has a rotating body which rotates under the motive power of a pressurised motor liquid flow. The motor has a high pressure region which receives the pressurised motor liquid flow, and a low pressure region through which the motor liquid flow leaves the motor. The pump also has a rotating body. The pump has a low pressure region which receives servo liquid flow to be pumped by its rotating body, and a high pressure region through which pressurised servo liquid flow leaves the pump. Each rotating body is mounted on a respective journal.

Abstract: An electronics enclosure for use in a gas turbine engine. The electronics enclosure including: a housing; one or more electronic devices, located within the housing, for use in controlling or monitoring the gas turbine engine; and a thermoelectric cooler configured, based upon a determination of an operational state of the gas turbine engine, to change between an active cooling mode, whereby the thermoelectric cooler is provided with power and thereby operated to cool the inside of the housing, and a passive mode, in which the thermoelectric cooler is not operated.

Abstract: A system for supplying lubricant to a component of a gas turbine engine is provided. The system includes a first fluid circuit having a first pump drivably couplable to a fan shaft of the engine for pumping a lubricant from a lubricant tank to the component. The first fluid circuit allows only a unidirectional flow of the lubricant from the lubricant tank to the component via the first pump. The system also includes a second fluid circuit having a second pump drivably couplable to the fan shaft for pumping the lubricant from the component to the lubricant tank. The second fluid circuit allows only a unidirectional flow of the lubricant from the component to the lubricant tank via the second pump. First and second ports of the two pumps operate as input ports or output ports depending on the rotation direction of the fan shaft.

Abstract: There is disclosed a containment assembly (500 for a gas turbine engine, comprising: an outer impact structure (56); an inner impact structure (54) disposed radially within the outer impact structure; wherein the outer impact structure and the inner impact structure axially extend between a fan portion (4) of the containment assembly corresponding to an axial location of a fan and a forward portion (2) of the containment assembly axially forward of the fan portion; and wherein at least one of the outer impact structure and the inner impact structure has a discontinuous profile between the forward portion and the fan portion so that there is a local reduction in compressive strength in the fan portion.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine. The system includes an oil tank located within a core of the engine, an oil filling reservoir located within a bifurcation that extends from the core of the engine, an oil tank filling pipe that leads from the oil filling reservoir to a tank filling port on the oil tank, and an oil pump. The system is configured so that the oil pump pumps oil from the oil filling reservoir through the oil tank filling pipe and the tank filling port to fill the oil tank. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: A gas turbine engine comprising a planetary gear train, and a core engine casing. The gear train has a ratio of greater than approximately 3.0, with an input to the gear train being operatively connected to the compressor section, and an output from the gear train being operatively connected to the fan. The core engine casing encloses the compressor section and the turbine section. The fan has a diameter F, and the core engine casing has a diameter C. The core engine casing diameter C varies along an axial length of the core engine casing, and a ratio (C/F) of the core engine casing diameter C to the fan diameter F is within the range 0.2<(C/F)<0.4, along an axial length of the core engine casing.

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.