Abstract: A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core.

Abstract: The present relates to combustor apparatus for a gas turbine engine comprising a bulkhead (34) an inner support ring (70) and an outer support ring (84) at end upstream end thereof. The bulkhead (34) has an inner surface (40), when in use is provided on an internal surface of a combustor (16) and exposed to combustion products and an outer surface (41) provided on an external surface of the combustor when in use; an inner mating feature to cooperate with the inner support ring (70); and an outer mating feature to cooperate with the outer support ring (84). In use, the inner and outer mating features prevent axial motion of the bulkhead relative to the inner (70) and outer (84) support rings.

Abstract: An aerofoil and an aerofoil assembly, in particular an aerofoil with improved stagnation zone cooling and an aerofoil assembly comprising such an aerofoil. The aerofoil is an aerofoil for a gas turbine engine comprising a pressure surface, a suction surface, a leading edge, a trailing edge, a stagnation zone located in the region of the leading edge, and an elongate channel running along the leading edge at the stagnation zone.

Abstract: There is provided an air intake system for providing air to a tip clearance control system. The air intake system comprises a ram-air intake having a scoop portion and a body portion. The body portion of the ram-air intake houses a heat exchanger.

Abstract: A complex cycle gas turbine engine (103) for an aircraft with hydrogen fuel supply. The gas turbine engine comprises, in fluid flow series, a core gas turbine (105) and a recuperator (701) for heating hydrogen fuel prior to combustion by turbine exhaust products.

Abstract: A gas turbine engine includes an engine core, a fan located upstream of the engine core, a nacelle surrounding the engine core and defining a bypass duct, and a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle. The engine core includes a compressor system, and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. An axial midpoint of a radially inner edge of the fan OGV is defined as the fan OGV root centrepoint. A fan OGV root position to fan diameter ratio of: an ? ? axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection ? ? and the ? ? fan ? ? OGV ? ? root ? ? centrepoint the ? ? fan ? ? diameter is equal to or less than 0.33.

Abstract: A clamping mechanism for machining a composite component. The clamping mechanism comprising two or more opposing clamping plates, each clamping plate comprising a clamping surface; and wherein each clamping surface is adaptable to follow the microstructural surface of the composite component being clamped. The clamping surface may comprise a plurality of deformable micropillars featuring a body that extends to a base on the clamping plates and a head that deforms to follow the microstructural surface of the composite component. The clamping surface in use may be an exact negative of the microstructural surface of the component clamping surface. Each clamping surface may comprise a disposable insert that replicates exclusively at the point of contact the microstructural surface of the component being clamped.

Abstract: A direct fuel injection system (206) is shown for injecting hydrogen fuel into a gas turbine combustor, the fuel injection system comprising a plurality of fuel injector blocks (1202). Each fuel injector block includes one or more air admission ducts (1603) for receiving air from a diffuser and an air outlet for delivering air into a mixing zone for combustion with fuel. Each fuel injector block also includes a fuel admission duct or aperture having a fuel inlet for receiving fuel (F) from a manifold, and a fuel outlet for delivering fuel into the mixing zone.

Abstract: A direct injection fuel system is shown for injecting hydrogen fuel into a gas turbine combustor. The fuel injection system includes a plurality of fuel injector blocks. Each fuel injector block includes a fuel admission duct having an inlet for receiving hydrogen fuel from a fuel supply, an outlet for delivering hydrogen fuel into the combustor and a central axis extending from said inlet to said outlet. Each fuel injector block also includes an air admission duct located around the periphery of the fuel admission duct, having an inlet for receiving air from a diffuser and an outlet for delivering air into the combustor for mixing with the hydrogen fuel.

Abstract: A composite storage tank system for gaseous hydrogen comprises a composite storage tank having composite wall enclosing a gas storage volume, the composite wall including a metal hydride element, or a metal element capable of forming a metal hydride in the presence of hydrogen, the system further comprising measuring apparatus arranged to measure an electrical characteristic of the metal hydride element or the metal element. The history of leakage of gaseous hydrogen from the tank, the current rate of leakage and the physical condition of the composite wall in the vicinity of the metal or metal hydride element may be inferred from a measurement of the electrical characteristic, without taking the tank out of service as is required in the case of known leaks tests such as a vacuum test, helium leak test or hydrogen sniffing test.

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

Abstract: A hydrogen fuel vaporiser (218) is shown for vaporising cryogenically-stored hydrogen fuel prior to injection into a gas turbine engine. The vaporiser (218) comprises a fuel offtake (301) configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit (217), a burner (303) configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit (217), and a heat exchanger configured and arranged to transfer heat produced by the burner to hydrogen fuel in the main fuel conduit (217).

Abstract: An operating condition monitor (100) for monitoring an operating condition of a transistor-based power converter (102), comprising: a sensing apparatus (106) configured to measure a turn-off transient energy of the power converter (102), a processor (108) in communication with the sensing apparatus (106) to receive the measurement of the turn-off transient energy, the processor being configured to: compare the measurement of the turn-off transient energy to a threshold; and issue an event signal based on the comparison to the threshold meeting a comparison criterion. A method (200, 200?) of monitoring an operating state of a transistor-based power converter is also disclosed.

Abstract: An aerofoil assembly includes a platform and one or more aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. The one or more aerofoils are disposed between the first edge and the second edge. Each of the one or more aerofoils has a leading edge proximal to the first edge and a trailing edge distal to the first edge. The platform defines one or more recesses disposed between the leading edge of each of the one or more aerofoils and the first edge.

Abstract: There is disclosed a roller interchange device 200 comprising: a support 202 for mounting on a composite material lay-up head 100; a plurality of roller mounts 208 for mounting respective rollers 312, 314. Each roller mount 208 is selectively moveable to a respective engaged position relative the support 202. The engaged position of each roller mount 208 corresponds to a respective roller being held at a compaction location relative the head when the support is mounted on the head.

Abstract: An engine core including a turbine, compressor, and a core shaft connecting the turbine and compressor; a fan located upstream of the engine core including a plurality of fan blades; and a gearbox. The gearbox is arranged to receive an input from the core shaft and to output drive to the fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and includes a sun gear, a plurality of planet gears, ring gear, and planet carrier to the mounted planet gears. The planet carrier has an effective linear torsional stiffness and the gearbox has a gear mesh stiffness between the planet gears and the ring gear. A carrier to ring mesh ratio of: the ? ? effective ? ? linear ? ? torsional stiffness ? ? of ? ? the ? ? planet ? ? carrier ? gear ? ? mesh ? ? stiffness ? ? between ? ? the ? ? planet ? ? gears ? ? and ? ? the ? ? ring ? ? gear is greater than or equal to 0.2.

Abstract: An aircraft engine having a turbine, the turbine having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height rhub, a maximum radial height rtip, and a radial extent between rhub and rtip, wherein, at every point along the radial extent of the blade, the blade has a modeshape value V1 for a blade first vibratory mode defined as V 1 = ? D 1 ? LE ? ? D 1 ? MC ? , and wherein, when the engine is operating between its maximum speed and 70% of that maximum speed, at least 80% of the radial extent of the blade has a modeshape value V1 from 0 to 1.5.

Abstract: An aircraft engine having a compressor, the compressor having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height rhub, a maximum radial height rtip, and a radial extent between rhub and rtip, wherein, at every point along the radial extent of the blade, the blade has a modeshape value V1 for a blade first vibratory mode defined as V 1 = ? D 1 ? LE ? ? D 1 ? MC ? , and wherein, when the engine is operating between its maximum speed and 70% of that maximum speed, at least 80% of the radial extent of the blade has a modeshape value V1 from 0 to 1.5.

Abstract: An aircraft engine having a fan, the fan having at least one flutter-resistant blade, the blade having a leading edge (LE), a trailing edge (TE), a midchord (MC), a minimum radial height rhub, a maximum radial height rtip, and a radial extent between rhub and rtip, wherein, at every point along the radial extent of the blade, the blade has a modeshape value V1 for a blade first vibratory mode defined as V 1 = ? D 1 ? LE ? ? D 1 ? MC ? , and wherein, when the engine is operating between its maximum speed and 70% of that maximum speed, at least 80% of the radial extent of the blade has a modeshape value V1 from 0 to 1.5.

Abstract: A system and a method for determining a high oil consumption in a gas turbine engine of an aircraft are provided. The method includes determining one or more engine and aircraft conditions. The one or more engine and aircraft conditions includes at least one of an oil quantity, an oil temperature, an oil pressure, an engine speed, an aircraft altitude, and an aircraft attitude. The method further includes determining a trend in oil conditions based on at least the one or more engine and aircraft conditions. The trend in oil conditions provides at least one of a rate of consumption of oil or a time duration of remaining oil. The method further includes determining the high oil consumption based on a comparison of the trend in oil conditions with a threshold or a comparison model.