Abstract: A fuel cell system includes a fuel pre-heater, a fuel cell stack and a cooling circuit which is arranged to implement a Rankine cycle and includes a condenser. The fuel pre-heater is arranged to heat a flow of liquid hydrogen provided to an input thereof to provide a flow of gaseous hydrogen. The system further includes a conveying apparatus arranged to convey the gaseous hydrogen to a fuel input of the fuel cell stack such that the gaseous hydrogen is in thermal contact with coolant fluid in the condenser. The size and mass of the condenser may thereby be reduced. The pre-heater is arranged to heat coolant fluid within the cooling circuit, thereby increasing the efficiency of the Rankine cycle.

Abstract: Scavenge systems for a gas turbine engine including a valve positioned at a first drainage outlet of an enclosure to selectively open and evacuate liquid from the cavity through the first drainage outlet in response to the valve being submerged in liquid are provided. Methods of reducing suction of air from an enclosure of a cavity of a gas turbine engine are further provided.

Abstract: The present disclosure provides an accessory power module for a gas turbine engine having a high-pressure spool and a low-pressure spool, the accessory power module comprising: a first electrical machine configured to couple to the low-pressure spool of the gas turbine engine; a casing housing the first electrical machine, the casing comprising at least one support element configured to support the first electrical machine within the casing, wherein the casing is configured for airflow therethrough; an accessory gearbox comprising an input coupling configured to couple to the high-pressure spool of the gas turbine engine, the accessory gearbox mounted to the casing; and a second electrical machine configured to couple to an output coupling of the accessory gearbox; wherein the casing is configured to be mounted along an inlet air flow path into the gas turbine engine.

Abstract: A method of operating a gas turbine engine, the gas turbine engine including an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a main gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system arranged to supply oil to lubricate the main gearbox; and a heat exchange system arranged to transfer heat between the oil and the fuel, the oil having an average temperature of at least 180° C. on entry to the heat exchange system at cruise conditions. The method includes controlling the heat exchange system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.

Abstract: A mechanism (30) and a method for inserting an elongate member (35) through an aperture of a body, along a longitudinal axis (35), the mechanism comprising a feed portion (42) comprising a feed actuator (43) configured to engage with and drive the elongate member along the longitudinal axis; and, a twist portion (44) comprising a twist actuator (82) configured to engage with the feed portion and rotate the elongate member about the longitudinal axis.

Abstract: In some examples, a system including a gas turbine engine, the engine including a high-pressure (HP) shaft; HP compressor; HP turbine, second shaft; second compressor; second turbine, the second turbine being coupled to the second compressor via the second shaft (e.g., LP shaft); and a generator coupled to the LP shaft. The generator is configured to generate electrical power from rotation of the LP shaft, and increase electrical power generated by the generator to increase a torque applied to the LP shaft by the generator, e.g., in combination with reduction in engine thrust, or in response to the detection of a stall and/or surge of the engine. The increase in torque applied to the second shaft is configured to increase a rate at which a rotational speed of the second shaft decreases, e.g., in combination with the reduction in engine thrust or during the stall/surge of the engine.

Abstract: A gas turbine engine control system and a method for limiting turbine overspeed in case of a shaft failure. An overspeed protection system detects a shaft failure of the engine. A variable stator vane mechanism adjusts stator vanes of a compressor of the engine in their rotational position, and is activated to move at least one row of the stator vanes into a closed position which blocks air flow through the compressor in case a shaft failure is detected. Bleed valves in the compressor have closed and open positions and divert in the opened position airflow through the compressor to a bypass channel. A detection system detects directly or indirectly the position of the stator vanes, and the control system sets the bleed valves to the open position when the detection system detects that the stator vanes have moved into the closed position.

Abstract: The invention relates to an assembly for a gas turbine engine which has a rotor disc, rotor blades and an annular blade retention plate, which adjoins the axially front side or the axially rear side of the rotor disc. The blade retention plate includes a radially outer region, which fixes the rotor blades axially, and a radially inner region, which is fixed axially by means of a securing ring in a recess on the axially front side or on the axially rear side of the rotor disc. Provision is made for the radially inner region of the blade retention plate to bear on the securing ring under axial preload. The invention relates further to a method for fastening a blade retention plate to a rotor disc of a gas turbine engine.

Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

Abstract: An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 300, 400, 500, 600, 700 configured to introduce a pressurised fluid into a core airflow A at a region downstream of the turbine 19, wherein a rearward movement of the turbine 19 or a shaft 26 in a shaft break event directly actuates the pressure equalisation apparatus 300, 400, 500, 600, 700 to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.

Abstract: A gas turbine engine configured to combust gaseous hydrogen fuel comprises a combustor comprising an annular combustion chamber outer casing surrounding an inner combustor case and a fuel manifold configured to provide gaseous fuel to a plurality of fuel injectors. The fuel manifold is formed integrally with the combustion chamber outer casing.

Abstract: An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 200 configured to introduce a pressurised fluid into a core airflow A at a region downstream of the turbine 19 in the event of a shaft break to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The pressure equalization apparatus comprises a sensor 216 configured to directly detect a shaft break event. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.

Abstract: An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 200, 300, 400, 500, 600, 700 configured to introduce a pressurised fluid into a core airflow A at a region directly downstream of the turbine 19 in the event of a shaft break to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.

Abstract: The invention regards an acoustic liner which comprises: a facing sheet that comprises holes, the facing sheet having a porosity and a facing sheet thickness, a backing sheet, and a plurality of cells arranged between the facing sheet and the backing sheet, the cells having a cell depth and each cell defining a cavity. A plurality of internal necks are provided that extend from the inner side of the facing sheet towards the backing sheet, each internal neck being located around a hole of the facing sheet, thereby extending the longitudinal length of the hole. The invention further regards a gas turbine engine with such acoustic liner.

Abstract: A friction bearing of a planetary gearbox has a first rotationally fixed component and a second component rotatably connected thereto. Oil in the region of an oil feed pocket of the first component is directed into the bearing clearance between the components. The oil is directed into the oil feed pocket via at least one first line that opens into the oil feed pocket. Oil is also directed via at least one second line into the oil feed pocket, the port region of said second line into the oil feed pocket in the circumferential direction of the bearing clearance and in the main rotation direction of the second component in relation to the first component and/or in the axial direction of the friction bearing being spaced apart from the port region of the first line into the oil feed pocket.

Abstract: A system for measuring loading on a test specimen. The system includes the test specimen arranged between a first loading bar and a second loading bar. The system further includes a first loading unit and a second loading unit configured to apply a first load and a second load to the first and second loading bars, respectively. The system further includes a first clamp and a second clamp configured to hold the first and second loading bars against the first and second loads, respectively. The system further includes a clamp actuating unit configured to selectively release at least the first clamp. The clamp actuating unit further includes a controller configured to electrically actuate at least one first electromechanical transducer from a retained state to a released state to release the first clamp, such that the first loading bar applies a first loading wave to the test specimen.

Abstract: The disclosure describes acoustic panels (such as for gas turbine engines) and techniques for forming acoustic panels. In some examples, the acoustic panel comprises a coversheet comprising an outer face, an inner face, and a plurality of apertures extend from the outer face to the inner face. The acoustic panel also comprises a 3D printed support layer comprising a polymer and defining a plurality of cell openings, wherein a first side of the support layer opposite the cell openings is coincident with the inner face of the coversheet, and wherein the 3D printed support layer is 3D printed directly onto the inner face of the coversheet.

Abstract: An additive manufacturing system may include an energy source, an optical system to modify and direct an energy beam from the energy source toward a component to form a melt pool, and a material delivery device to deliver material to the melt pool. The optical system may form an annular energy beam, direct the annular energy beam toward the component, receive at least a portion of thermal emissions produced by the annular energy beam and the melt pool, and direct the portion of the thermal emissions toward an imaging device, which may be used to control the energy source.

Abstract: A computer-implemented method including: receiving a first data set including a plurality of values for a plurality of features; removing one or more features and associated values from the first data set to generate a second data set; determining feature importance of at least a subset of the features of the second data set using multiple-evaluation criteria; and performing an action using at least one feature of the second data set and the determined feature importance.