Abstract: There is disclosed a method for spatiotemporal registration between a first image stream and a second image stream. A first registration frame set derived from the first image stream and a second registration frame set derived from the second image stream are defined. A spatial correlation analysis is conducted based on the first registration frame set and the second registration frame set to determine a spatial registration offset. The first registration frame set and the second registration frame set are spatially registered based on the spatial registration offset. A temporal correlation analysis is conducted based on the first registration frame set and the second registration frame set after spatial registration to determine a temporal registration offset. The first image stream and the second image stream are spatially and temporally registered based on the spatial registration offset and the temporal registration offset.

Abstract: A method of detecting a shaft break in a turbine engine is provided. The shaft connects a torque-exerting turbine of the engine to a torque-receiving unit of the engine. The method includes: receiving repeated measurements of the rotational frequency of the torque-receiving unit and the rotational frequency measurements of the turbine; determining from the measurements an amount of shaft twist between the turbine and the torque-receiving unit; comparing the amount of twist with a threshold twist value; and signalling a shaft break when the amount of twist exceeds the threshold twist value and a pre-condition is met. The pre-condition includes determining from the measurements that the turbine rotational speed is greater than the simultaneous torque-receiving unit rotational speed by a threshold speed value.

Abstract: A combustion staging system includes a splitting unit which receives a metered fuel flow and controllably splits the received fuel flow into pilot and mains flows. Pilot and mains fuel manifolds distribute fuel from the splitting unit to the pilot and mains stages. The splitting unit selects and deselects pilot-only operation. Both pilot and mains manifolds are selectable for pilot and mains operation. A cooling flow recirculation line has a delivery section arranged to provide a cooling flow of fuel to the mains manifold when it is deselected during pilot-only operation. Cooling flow enters the delivery section from a high pressure fuel zone of the engine and exits the return section to a low pressure fuel zone of the engine. A controller adjusts the splitting unit during pilot-only operation to partially select the mains manifold thereby increasing the pressure in the mains manifold to meet a target fuel pressure therein.

Abstract: A method is provided of forming a neural network for detecting surface defects in aircraft engine components. The method includes: providing (i) a pre-trained deep learning network and (ii) a learning machine network; providing a set of pixelated training images of aircraft engine components exhibiting examples of different classes of surface defect; training the trainable weights of the learning machine network on the set of training images.

Abstract: The disclosure relates to a ceramic matrix composite (CMC) aerofoil. Example embodiments include an aerofoil comprising first and second tubular CMC cores (302, 303) extending along a longitudinal axis of the aerofoil; and an outer CMC layer surrounding the first and second tubular CMC cores (302, 303) and defining an outer shape of the aerofoil having leading and trailing edges, wherein fibres within a wall of the second tubular CMC core extend to the trailing edge of the aerofoil.

Abstract: An article 6 made from a porous material, such as a ceramic material, is dipped into a liquid impregnating material 12, such as molten urea in a bath 14. After withdrawal from the bath 14, excess impregnating material 12 is removed from the article 6 by means of an air knife 22. The air knife 22 is directed obliquely into an enclosure 18, which is maintained at a temperature above the melting point of urea. The article is passed into the enclosure 18 through the flow from a further air knife 20 which creates a separation airflow separating the interior of the enclosure 18 from the ambient surroundings. The article is carried by a fixture 6 to be conveyed along the processing line on a conveyor 2.

Abstract: The present disclosure relates to the inspection and repair/maintenance of an item such as a gas turbine engine. Apparatus (110) is provided for connecting apertures (124,142) in an item to be inspected. The apparatus (110) comprises a tubular guide (130) with first and second ends (134,136) and a hollow interior (138) for receiving an inspection tool, and a housing fixture (120) for mounting the tool to an outer shell (126) of the item. The apparatus (110) provides an articulated joint between the tubular guide (130) and the housing fixture (120). A cooling system may also be included.

Abstract: A component for a gas turbine engine, comprising: a first wall defining an exterior surface of the component; a second wall, arranged such that a coolant channel is defined by the space between the first and second walls; and a plurality of apertures provided through the first wall to connect the coolant channel to the exterior surface of the component; wherein adjacent at least one aperture the coolant channel comprises a flow modifier, configured to locally change the pressure of the coolant flowing in the coolant channel in the region of the aperture relative to a region of the coolant channel adjacent another aperture.

Abstract: A gas turbine engine comprises a gearbox comprises 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. The planet gear carrier comprising a first ring, a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring. Each planet gear is rotatably mounted on a respective one of the axles and the axles are arranged at a first radius. At least one of the first ring and the second ring comprises a metal matrix composite material ring and the metal matrix composite ring comprises a ring of reinforcing fibres and the ring of reinforcing fibres having a second radius greater than the first radius.

Abstract: A method of attaching a projection to a thin walled component, the thin walled component having a first surface and an opposite second surface, the method comprises the steps of: (i) providing a thin walled component; (ii) identifying a first position on the first surface of the thin walled component; (iii) building a hollow thin walled lattice structure on the second surface at a second position using a liquid metal deposition process, the second position corresponding to the first position; and (iv) building a projection on the first surface at the first position using a liquid metal deposition process.

Abstract: A nanocrystalline bainitic steel consisting of, by weight percentage: 0.3% to 0.6% carbon; 9.0% to 20.0% nickel; up to 10% cobalt; 1.0% to 4.5% aluminium; up to 0.5% molybdenum; up to 0.5% manganese; up to 0.5% tungsten; up to 3.0% chromium; and the balance being iron and impurities.

Abstract: The present disclosure relates to techniques for stress relief in additive layer manufacturing (ALM). Example embodiments include a method for additive layer manufacturing of a metallic component, comprising the steps of: providing a substrate (20); depositing a first layer (22) of material on the substrate (20); depositing a plurality of second layers of material on the first layer (22) to form the metallic component (21), wherein the first layer (22) forms a stress relieving layer between the plurality of second layers and the substrate (20), the stress relieving layer having a lower shear stiffness compared to the metallic component (21).

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: Disclosed herein is a method of manufacturing a double walled section of an aerofoil (701) of a gas turbine engine (10), the method comprising: forming a plurality of columns (402) on an outer surface of a first structure; forming a plurality of columns (402) on a surface of each of a plurality of parts of a second structure; and forming a double walled section of an aerofoil (701) by attaching the ends of columns (402) on each part of the second structure to the ends of columns (402) on the first structure such that the first structure is an inner wall (401) of the section of the aerofoil (701) and the second structure is an outer wall (501) of the section of the aerofoil (701).

Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: A turbofan engine includes an outer fixed structure, the downstream terminal end of which terminates in a terminal plane substantially normal to the longitudinal axis of the engine. The radius R of the internal surface of the outer fixed structure in the terminal plane is a function R(?) of azimuthal position ? such that the radius has a constant value R0 within a first azimuthal interval of 180°, a value greater than R0 at any azimuthal position within a second and fourth azimuthal intervals and a value less than R0 at any azimuthal position within a third azimuthal interval, the azimuthal intervals forming a total interval of 360° and the fourth interval being contiguous with the first. The engine may be mounted closer to the airframe of an aircraft than a turbofan engine having an outer fixed structure which is axisymmetric in its downstream terminal plane.

Abstract: A gas turbine engine (10) comprising: an engine core (11), comprising a compressor (14, 15); an outer casing (25A) separating the engine core (11) from a bypass airflow; a compressor bleed valve (50) in communication with the compressor (14, 15) and configured to release bleed air from the compressor (14, 15); a bleed air duct (51) connected to the compressor bleed valve (50) and configured to eject the bleed air released by the compressor bleed valve (50) into an airflow at a location radially inward of the outer casing (25A).

Abstract: A shaft assembly for a gas turbine engine is provided. The shaft assembly comprises: a first shaft having an outer surface; a first coupling ring disposed around the outer surface of the first shaft, an inner surface of the first coupling ring being coupled to the outer surface of the first shaft; a second shaft having an inner surface; and a second coupling ring disposed around the inner surface of the second shaft, an outer surface of the second coupling ring being coupled to the inner surface of the second shaft, wherein an outer surface of the first coupling ring is configured to mate with an inner surface of the second coupling ring, such that concentricity of the first and second shafts is maintained at the shaft assembly by virtue of the mating of the first and second coupling rings. Methods of assembling and re-assembling a shaft assembly are also provided.

Abstract: Welding apparatus, for example for a workpiece, comprises resistance welding electrode wheels which form a weld seam between components of the workpiece. The formed weld is monitored by an ultrasonic probe which is situated behind the rollers with reference to the weld direction. The ultrasonic probe comprises a linear phased array of piezoelectric elements. The probe enables real-time monitoring of the welding process, with the facility to adjust the process to correct deviations.

Abstract: A composite aerofoil, the aerofoil having a leading edge, a trailing edge a pressure side and a suction side and comprising an outer erosion protection layer along one or both of the pressure side or suction side, a structural core having a plurality of resin impregnated plies of unidirectional fibers, and between the outer erosion protection layer and the structural core a woven composite impregnated with a resin having a modulus of elasticity greater lower than that of the resin impregnating the plies in the structural core.