Abstract: An electrical fault detector is shown for installation in electrical network (101) of the type comprising a first voltage source (104) and a second voltage source (107), each of which have a respective positive rail (105,108) connected by a positive concentrator (110) and a respective negative rail (106,109) connected by a negative concentrator (111). The detector comprises an inductor (112) for location in one of: the positive concentrator between the connections of the positive rails thereto, and the negative concentrator between the connections of the negative rails to thereto. The detector also comprises a fault identification device (113) configured to monitor the voltage across the inductor, and generate a fault signal in response to the voltage across the inductor exceeding a threshold.

Abstract: An elongate compressible seal is provided for sealing between a first component and a second component, the seal comprising a plurality of elongate members each attached at a first end to an outer surface of the seal for aiding positioning of the seal.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial location ? ? of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage ? ground ? ? plane ? ? to ? ? wing ? ? distance and a quasi-non-dimensional mass flow rate Q defined as: Q = W ? T ? ? 0 P ? ? 0 · A flow where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; and Aflow is the flow area of the fan face in m2, and wherein a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 to 0.01.

Abstract: The present disclosure concerns an electrical interconnect system for a vehicle such as an aircraft or marine/submarine vessel.

Abstract: A gas turbine engine includes an engine core; a fan located upstream of engine core and having a fan diameter; a nacelle surrounding engine core and defining a bypass duct between engine core and nacelle; and a fan OGV extending radially across bypass duct, fan OGV having a radially inner and outer edge, wherein the radially inner edge's axial midpoint is the fan OGV root centrepoint. The engine core includes a compressor system encompassing a first, lower pressure, compressor, and second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of outer core casing at an axial position of first flange connection, the first flange connection having a first flange radius, wherein the first flange connection is downstream of an axial position.

Abstract: A turbomachinery control system for controlling supercritical working fluid turbomachinery. The control system includes a light emitter to project light through working fluid of the turbomachinery toward a primary light detector provided within a line of sight to the emitter. The system further includes one or more secondary light detectors spaced from the line of sight, and a controller determining one or both of an intensity of light detected by the primary detector relative to the detected light intensity by the secondary detector, and wavelength of light detected by the primary detector relative to wavelength of light detected by the secondary detector. The controller determines the working fluid proximity of the critical point based on one or both of the determined relative intensity and determined relative wavelength, and controlling an actuator to control turbomachinery inlet or outlet conditions in accordance with the working fluid determined proximity of the critical point.

Abstract: A method of providing datum features on a ceramic article including the steps of providing a ceramic member having a plurality of substantially parallel holes extending into the ceramic member. Providing a plurality of tubular mounts, the cross-sectional area of each tubular mount being less than the cross-sectional area of the corresponding hole. Applying adhesive to the exterior surface of each one of the tubular mounts and inserting an adhesive covered tubular mount into each one of the holes. Providing a fixture member having a plurality of substantially parallel projections extending from predetermined positions. Positioning the ceramic member on the fixture member such that each projection locates in one of the tubular mounts and fixes the tubular mounts at the predetermined positions. The ceramic article may be removed from the fixture member or located in a machine tool using the fixture member and machining the ceramic article.

Abstract: A gas turbine engine assembly adapted to separate a high pressure zone from a low pressure zone includes a seal assembly configured to block gasses from passing through the interface of two adjacent components. The seal assembly includes a rod.

Abstract: A bush having a neck portion (33) for inserting into an aperture of one or both of two components to be fastened together, a radially extending shoulder portion (34) having a rear face (34a) adjacent the neck portion (33) an opposite front face (34b) and a radially outer face (34c) extending between the rear face (34a) and the front face (34b), one or more ventilation slots (36) provided in the rear face extending from a junction between the rear face (34a) and the neck portion (33) whereby, in use, to provide an escape route for air caught between the bush and a component.

Abstract: A rotor of a turbomachine includes a plurality of rotor blades, which each have a blade leading edge and a tip, the rotor blades forming a first group and at least one further group of rotor blades of different construction. It is provided that the rotor blades of the first group are nominal rotor blades, the rotor blades of the at least one further group each have a material cutout in the transition from the blade leading edge to the tip, the material cutout not being present in the case of the rotor blades of the first group, and the rotor blades of the first group and of the at least one further group form a non-periodic sequence in a circumferential direction.

Abstract: An aircraft propulsion system comprises first and second co-axial propulsors, one of the first and second propulsor being positioned forward of the other propulsor. A first electric motor is configured to drive the first propulsor, and a second electric motor is configured to drive the second propulsor. The first electric motor comprising a rotor radially inwardly of the stator, and the second electric motor comprises a rotor radially outwardly of the stator. The stator of the first electric motor is mounted to the stator of the second electric motor.

Abstract: An aircraft propulsion system comprises first and second propulsors. An epicyclic reduction gearbox is coupled to an input shaft. The reduction gearbox comprises a sun gear coupled to the input shaft, a plurality of planet gears which mesh with the sun gear and are mounted to a planet carrier, and a ring gear which meshes with the plurality of planet gears. The first propulsor is located upstream of the second propulsor The first propulsor is coupled to the sun gear, and the second propulsor is coupled to one of the planet carrier and the ring gear.

Abstract: An aircraft propulsion system comprises a vertical lift propulsor mounted within a wing of the aircraft. The vertical propulsor comprises an electric motor comprising a rotor mounted to a plurality of propulsor blades, the motor further comprising a stator. The rotor is provided radially outwardly of the stator.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The space to chord ratio of the aerodynamic vanes that are closest to the suction surface of a strut-vane is higher than the space to chord ratio of aerodynamic vanes that are closest to a pressure surface of the strut-vane. The arrangement allows the duct 100 to be axially short.

Abstract: A method of measuring the perimetric width of a hollow component. The hollow component comprises a first panel and a second panel, with the first and second panels being bonded to one another around their respective perimeters, and a perimetric width being defined as the width of the bonded portion joining the first panel to the second panel.

Abstract: A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

Abstract: Methods and systems automatically detect, classify and/or mitigate sensor errors using partial qualitative and quantitative knowledge of the subsystems. In various examples, sensor fault detection is performed with a custom designed representation scheme covering causality, correlation, system of equality and inequalities, and an associated logic. The logic is described by a set of algorithmic steps to iteratively assign trustworthiness level of sensors. Sensor fault classification is performed by combining mathematical and statistical techniques that can be utilized to expose bias, drift, multiplicative calibration error, precision degradation and spike error. Sensor fault mitigation is also performed on identified bias, drift, multiplicative calibration error, precision degradation and spike error.

Abstract: A method of forming a fibre composite component (10), the method comprising: providing a plurality of first fibres (102), the first fibres (102) being arranged in a plurality of layers (104a-e) with an uncured thermoset resin matrix (106) being provided therebetween. The resin matrix (106) comprises a plurality of second electrically conductive fibres (108) suspended therein having a shorter length than the first fibres (102). The method comprises then applying an energy field to the resin matrix (106) to thereby align the second fibres (108) such that they provide an electrically conductive path between adjacent layers (104a-e) of first fibres (102) to form an electrically conductive fibre network comprising the first and second fibres (102, 108). The method comprises then inducing an electric current through the electrically conductive fibre network to thereby heat and cure the resin matrix (106).

Abstract: A turbine wheel for use in a gas turbine engine having a plurality of blades attached to a rotor disk. The blades each fit within slots formed in the rotor disk to couple the blades to the rotor disk. Platform segments are arranged about each of the blades.

Abstract: A composite component comprising fibres embedded in a matrix material. A plurality of pins extend through the matrix material. Each pin of the plurality of pins comprises a hollow bore such that the bore of each pin defines a plurality of holes in the composite component.