Abstract: An aircraft propulsion system (10) comprises at least first and second electrical generators (15a, 15b), each being configured to provide electrical power to a respective first and second AC electrical network (16a, 16b). The system (10) further comprises at least first and second AC electrical motors (19a, 19b) directly electrically coupled to a respective AC network (16a, 16b) and coupled to a respective propulsor (4), and a DC electrical network electrically coupled to the first and second AC networks (16a, 16b) via respective first and second AC to DC converters (17a, 17b), and to a further electrical motor 19c), the further electrical motor (19c) being coupled to a propulsor (4).

Abstract: A fluid atomizer and methods of atomizing fluids are disclosed. The fluid atomizer may comprise an inner member and one or more outer members. The inner member defines an interior conduit for providing a first-fluid flowpath from a supply end of the atomizer to a discharge end of the atomizer along a central axis. The one or more outer members are positioned radially outward of the inner member from the central axis. The inner and outer members define a second-fluid flowpath extending from a second-fluid supply conduit to a second-fluid discharge plenum. The second-fluid flowpath comprises a tangential conduit spiraling along the axis from the second-fluid supply conduit to a terminal end; an annulus downstream from and in fluid communication with the tangential conduit; and a second-fluid discharge plenum downstream from and in fluid communication with the annulus.

Abstract: A combustion staging system is provided for fuel injectors of a multi-stage combustor of a gas turbine engine. The system has a splitting unit which receives a metered total fuel flow and controllably splits the metered total fuel flow into out-going pilot and mains fuel flows to perform pilot-only and pilot-and-mains staging control of the combustor. The system further has pilot and mains fuel manifolds which respectively receive the pilot and mains fuel flows. The system further has a plurality of mains flow scheduling valves which distribute the mains fuel flow from the mains fuel manifold to mains discharge orifices of respective injectors of the combustor. The system further has servo line which extends to the mains flow scheduling valves, the servo line being controllably fillable with pressurised fuel to change the pressure in the servo line relative to the pressure in the mains fuel manifold.

Abstract: Apparatus for a gas turbine engine, the apparatus comprising: a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing; a plurality of guide vanes extending radially from the outer wall of the core engine casing; a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; and an accessory gear box positioned within the second cavity of the torque box.

Abstract: Energy conversion apparatus comprising: an electrical machine including a rotor arranged to rotate about an axis, the rotor defining a first cavity therein; power electronic circuitry arranged at least partially around the axis and defining a second cavity therein, the power electronic circuitry being positioned adjacent to the electrical machine; a heat pipe positioned within the first cavity of the rotor and within the second cavity defined by the power electronic circuitry, the heat pipe being arranged to receive thermal energy from the rotor; and a cooling arrangement positioned at least partially within the second cavity and arranged to receive thermal energy from the power electronic circuitry and from the heat pipe.

Abstract: A method of calibrating an engine core of a gas turbine engine, wherein the engine core includes a turbine, combustion equipment, a compressor, and a core shaft connecting the turbine to the compressor, the core shaft arranged to drive a propulsive fan of the gas turbine engine, the method including: providing a resistance load on the core shaft, the resistance load arranged to replicate the load of a propulsive fan; driving the engine core; measuring a performance parameter or the engine core; measuring a thrust generated by the engine core; and determining power rating data of the engine core, providing a correlation between the performance parameter and the thrust.

Abstract: There is described a device for controlling an amount of current within a power distribution network by manipulating the amount of magnetic flux in the device and thus the impedance experienced by the power distribution network across the device. This is achieved by winding a plurality of coils about a magnetically permeable core and by providing the device with a magnetically permeable bridge element that is movable between a fully-open position at which the net magnetic flux generated in the core by alternating currents in each coil is zero, and a fully-closed position at which a net magnetic flux is present in the core.

Abstract: A combustion chamber arrangement includes an annular combustion chamber, a plurality of lean burn fuel injectors and a stage of turbine nozzle guide vanes. The stage of turbine nozzle guide vanes is arranged at the downstream end of the annular combustion chamber and includes a plurality of circumferentially spaced vanes extending between and being secured to an annular inner wall and an annular outer wall. The annular inner wall has an outer surface which has a plurality of outer surface portions and each portion is positioned between a pair of circumferentially adjacent vanes and the annular outer wall has an inner surface which has a plurality of inner surface portions and each portion is positioned between a pair of circumferentially adjacent vanes.

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: 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: An engine for an aircraft has 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 for an aircraft is arranged to receive an input from a core shaft and to output drive to a fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and has a sun gear, a plurality of planet gears, a ring gear, and a planet carrier having a plurality of pins, each pin being arranged to have a planet gear of the plurality of planet gears mounted thereon. A ratio of planet carrier torsional stiffness to pin stiffness is within a specified range.

Abstract: An electrical power generator system has: an input shaft for receiving a rotary input drive; plural output shafts connected by respective gear shifting arrangements to the input shaft thereby providing different gear ratios between the input shaft and each output shaft; and plural electrical generators powered by rotation of the respective output shafts, electrical power outputs of the generators being combined to supply a total power to a load. The input shaft operates over a range of rotation frequencies, and the gear shifting arrangements are configured to shift the gear ratios between the input shaft and the output shafts such that the output shafts operate over a narrower range of rotation frequencies.

Abstract: There is disclosed a turbine comprising a stator vane ring 122 downstream of a rotor ring 116 of shrouded rotor blades 118. In vanes of the stator vane ring, an upstream portion of a mean camber line 136 diverges from a profile of the mean camber line at a half-span location, particularly to receive a tip leakage flow over the shrouded rotor blades 118.

Abstract: A planet carrier for an epicyclic gearbox, the planet carrier comprising: a single piece carrier body arranged to support a plurality of gears of the epicyclic gearbox; and an opening in a wall of the carrier body; wherein: in a cross-section of the carrier body in a plane that is orthogonal to the longitudinal axis of the carrier body and at a longitudinal position along the carrier body that does not comprise the opening, the carrier body is substantially annular; and the opening is arranged such that the gears that the carrier body is arranged to support can be passed entirely through the opening.

Abstract: The blades for a rotor of a gas turbine engine are all manufactured to the same design. However, manufacturing tolerances mean that in practice each individual blade is different to the others. It is proposed to arrange the blades around the circumference of the rotor in a manner that limits excessive stress being induced in the blades due to differences in the vibration response between a given blade and its two neighbouring blades.

Abstract: A gas turbine engine for an aircraft includes: a flow path boundary, which delimits the flow path through the engine radially on the outside, and a lining, which lines the flow path boundary on the inside, at least along an axial section. Here, the lining includes a plurality of panels, which, in the circumferential direction of the flow path boundary, adjoin each other and which together line a circumferential area of 360°, wherein each panel has two end faces, which each adjoin an end face of an adjacent panel. The panels are of beveled design at their end faces, such that two mutually adjoining panels form a V-shaped gap between them, the minimum clearance of which is realized at the inside of the panels. The panels can be sound-absorbing panels. Also disclosed is a panel for a gas turbine engine.

Abstract: A gas turbine engine comprising: a combustor configured to initiate combustion; and a turbine comprising a stator vane ring defining a plurality of passageways between adjacent vanes; wherein at least one of the passageways is provided with a restrictor which defines a temporary gas washed surface for the stator vane ring and is configured to be ablated upon initiation of combustion to reveal an operational gas washed surface of the stator vane ring. A method of starting a gas turbine engine is also described.

Abstract: There is disclosed a vane for a gas turbine engine, the vane comprising a three-dimensional weave of composite material. The vane is hollow to define a hollow core, wherein the vane has a maximum thickness (T), and a maximum wall thickness (WT) between an inner surface and an outer surface in a direction normal to the outer surface, wherein WT?0.1T. Also disclosed is a method of manufacturing a composite vane for a gas turbine engine, the method comprising: forming a preform for the vane comprising a three-dimensional weave of composite material; and curing the preform. The preform has an internal hollow region, and the cured vane has a maximum thickness (T), and a maximum wall thickness (WT) between an inner surface and an outer surface in a direction normal to the outer surface, wherein WT?0.1T.

Abstract: A gas turbine engine casing is described as having a cowl door hinged to a casing support structure by at least one hinge. The cowl door is openable outwardly from the casing to expose a casing interior. The hinge is located above a longitudinal axis of the casing and comprises a pivoting linkage arranged such that, upon actuation between closed and open cowl door conditions, the pivoting linkage moves an upper portion of the cowl door downwards towards the longitudinal axis.

Abstract: A clearance control arrangement (26) for a rotor (28) comprising a segment assembly (33) spaced radially inwards from a casing (32) and defining a radial clearance (42) between the rotor (28) and the segment assembly (33). A front carrier (35a) and a rear carrier (35b) supporting the segment assembly (33). A heat transfer cavity (48) formed between the carriers (35a, 35b), the segment assembly (33) and the casing (32). A settling chamber (60) formed adjacent to the casing (32) and upstream of the heat transfer cavity (48). The clearance control arrangement (26) is configured to receive air into the settling chamber (60) and thence to deliver it to the heat transfer cavity (48).