Abstract: A composite gas storage tank includes a composite wall defining a main storage volume and a hollow conduit portion communicating with and extending away from the main storage volume. A hollow coupling element has a length portion which is partially embedded within and extends substantially parallel to the hollow conduit portion of the composite wall. The hollow conduit portion and the coupling element provide communication between the main storage volume and the tank exterior. A leakage path around the hollow coupling element is significantly longer than those of tanks of the prior art and reduces the rate at which hydrogen leaks from the tank. A carbon fibre filament winding extends over the length of the hollow conduit portion and provides a radially inwardly-directed force biasing the carbon fibre material of the hollow conduit portion into contact with the coupling element, further reducing the leakage rate on the leakage path.

Abstract: A vane for a gas turbine engine, the vane including a platform with an airfoil extending radially from the upper surface of the platform. The platform includes a joint portion which includes a circumferentially extending flange and a recessed surface both formed on either the upper or lower surface of the platform. The flange and the recessed surface extend from opposing circumferential edges of the joint portion and each include a substantially radially-extending through hole.

Abstract: There is provided an air pressurisation system comprising: a blower compressor coupled to a spool of a gas turbine engine and configured to receive inlet air from a bypass duct of the gas turbine engine; an air treatment apparatus to mix air from the air cycle line with air from the bypass line to control a temperature of air for discharge to an airframe system; and a core bleed line configured to provide a second inlet flow of air from a compressor of the gas turbine engine to the air treatment apparatus in an augmented air supply mode of the air pressurisation system; wherein the core bleed line is configured to provide the second inlet flow of air directly to the bypass line or mix the second inlet flow of air with air from the blower compressor upstream of the bypass line.

Abstract: A system for use with a component of a power generation apparatus includes a lens arrangement disposed proximal to the component. The system further includes a variable focal length structure spaced apart from the lens arrangement and optically coupled with the lens arrangement. The variable focal length structure has a focal plane that is adjustable. The system further includes a controller communicably coupled with the variable focal length structure. The controller is configured to control the variable focal length structure to adjust the focal plane of the variable focal length structure, such that at least one region of interest of the component is in focus.

Abstract: A power electronics converter comprising: first and second input terminals; first and second DC output terminals; a branch comprising first and second semiconductor switches connected in series between the first and second DC output terminals, the first input terminal connected to a node between the first and second semiconductor switches; a DC link capacitor connected between the first and second DC output terminals; and a reverse biased DC link diode connected across the DC link capacitor.

Abstract: A method for improving the corrosion and fatigue crack resistance of an article that has a surface. The method involves: providing a dry powder having one or more anti-corrosion materials; pressurising the dry powder to 1 to 5 MPa in an inert gas; pre-heating the surface of the article to a temperature of 50 to 900° C. and cold-spraying the surface of the article with the dry powder at a particle velocity of 300 to 1200 m/s so that a corrosion and crack-resistant coating is formed on the surface of the article and the surface is modified to an average depth of 10 to 100 ?m.

Abstract: A gas turbine engine includes a core engine casing and a bypass duct defined between a nacelle and the core engine casing. The gas turbine engine further includes a plurality of diverter fences pivotally coupled to the core engine casing. Each diverter fence is pivotable relative to the core engine casing about a pivot axis, which is circumferentially and obliquely inclined with respect to a principal rotational axis. Each diverter fence is configured to move between a first position in which an outboard edge is disposed adjacent to a casing outer surface, and a second position in which the outboard edge is radially spaced apart from the casing outer surface, such that each diverter fence radially extends outwards from the casing outer surface into the bypass duct.

Abstract: In some examples, a method comprising: controlling a first ratio of a first liquid to a second liquid to form a first suspension comprising a powder and a first carrier liquid composition comprising at least one of the first liquid or the second liquid; directing the first suspension comprising the first carrier liquid and the powder to a plume of a thermal spray device; forming a first portion of a coating comprising the powder on a substrate from the first suspension; controlling a second ratio of the first liquid to the second liquid to form a second suspension comprising a second carrier liquid composition and the powder; directing the second suspension comprising the second carrier liquid composition and the powder to the plume of the thermal spray device; and forming a second portion of the coating comprising the powder on the substrate from the second suspension.

Abstract: A gas turbine engine for an aircraft. The gas turbine comprises a staged combustion system having pilot injectors and main injectors, a fuel metering system configured to control fuel flow to the pilot injectors and the main injectors, and a fuel system controller. The controller is configured to identify an atmospheric condition, determine a ratio of pilot fuel flow rate for the pilot injectors to main fuel flow rate for the main injectors in response to the atmospheric condition, and inject fuel by the pilot injectors and the main injectors in accordance with said ratio to control an index of soot emissions caused by combustion of fuel therein.

Abstract: A blade for use in a gas turbine engine, the blade comprising a blade portion and a fir tree root portion, the blade portion and the root portion having a connected passage for allowing cooling air to flow within the blade, the fir tree root portion having an air intake on its leading edge, the air intake allowing cooling air to enter the cooling passage and wherein the fir tree root portion comprises a plurality of projections, including at least a base projection and a top projection; and wherein the air inlet located in the base projection of the fir tree root portion and wherein the air inlet comprises at least 50% of the face of the base projection of the fir tree root portion.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

Abstract: A method of optimising the performance of combustion equipment of a gas turbine engine includes providing a fuel flow into the combustion equipment via a plurality of fuel injectors circumferentially disposed about a principal rotational axis at a plurality of injector positions; determining a plurality of temperatures of combustion gases at a plurality of circumferential positions downstream of the combustion equipment using a plurality of temperature measurement devices; ranking the plurality of circumferential positions based on the plurality of temperatures of the combustion gases determined using the plurality of temperature measurement devices; and repositioning at least some of the plurality of fuel injectors between the plurality of injector positions based at least on the ranking of the circumferential positions.

Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: 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 that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

Abstract: A combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

Abstract: A bladed disc system for a turbine engine having a disk portion and a plurality of blade portions which are associated with a stator section and an intercavity sealing portion, disc portion shaped such that blade portions are able to fit within firtree slot in disc portion, blade portion having aerofoil section and root section, aerofoil section having portion shaped such that they extend proximate to intercavity sealing portion, disc portion extending from portion that connects with drum to outer edge at which blade portions are connected with disc portion having width transition region in which thickness of disc increases from point at which disc connects to drum to outer edge at which it holds blade portions, and wherein width transition region has curved width transition region with radius r, and an overhanging portion which extends into the intercavity spacing between the width transition region and the intercavity sealing portion.

Abstract: A turbofan gas turbine engine having a bypass duct, and a bypass airflow measurement system. The bypass airflow measurement system comprises: at least one acoustic transmitter configured to transmit an acoustic waveform across the bypass duct of the gas turbine engine though which a bypass airflow passes to at least one acoustic receiver; where the at least one acoustic transmitter and the at least one acoustic receiver are located on an axial plane that is substantially perpendicular to the bypass flow. A method of measuring bypass airflow properties of a turbofan gas turbine engine is also described.

Abstract: A method of operating an aircraft including a gas turbine engine and a plurality of fuel tanks arranged to provide fuel to the gas turbine engine, where at least two of the fuel tanks contain fuels with different fuel characteristics. The method includes obtaining a flight profile for a flight of the aircraft; and determining a fuelling schedule for the flight based on the flight profile and the fuel characteristics. The fuelling schedule governs the variation with time of how much fuel is drawn from each tank. Fuel input to the gas turbine engine may then be controlled in operation in accordance with the fuelling schedule.

Abstract: A fuel spray nozzle includes a fuel circuit having in series a gallery, circumferentially spaced passages arranged in a row around the nozzle, and an annular spin chamber. Each passage has an inlet for receiving a respective portion of the fuel from the gallery and a metering orifice for discharging its portion of the fuel. The passages are configured such that, when the flow of liquid fuel to the inlet port is shut off, a respective differential static pressure develops across stagnant liquid fuel remaining between the inlet and the metering orifice of each passage, and the passages are further configured such that one or more selected passages develop a different differential static pressure to the remaining passages causing a flow of purging air to enter the gallery from the combustor through the selected passages and exit through the remaining passages, thereby purging the gallery and the passages of fuel.

Abstract: A computer-implemented method controls input of at least a portion of a first training data set into a first machine learning algorithm. The first training data set includes data quantifying damage to a first compressor and data quantifying a first operating parameter of the first compressor. The first machine learning algorithm is executed, and data quantifying the first operating parameter is received as an output of the first machine learning algorithm. The first machine learning algorithm is trained using the received data output from the first machine learning algorithm and data quantifying the first operating parameter of the first compressor. The trained first machine learning algorithm is configured to enable determination of operability of a second compressor of a gas turbine engine.

Abstract: A gas turbine engine comprises a fan; an engine core comprising a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox; an Engine Section Stator (ESS) comprising a plurality of ESS vanes with an external surface washed by the core airflow; an ESS heating system adapted to heat the ESS vanes, and a temperature sensor adapted to detect the temperature of the external surface of the ESS vanes and send a signal to the ESS heating system when said temperature is below a reference temperature. Upon detection and/or inference of ice crystal conditions and receiving from the temperature sensor the signal that the temperature is below the reference temperature, the ESS heating system is activated to heat at least a portion of the external surface of the ESS vanes and promote melting and adhering of ice crystals thereto.