Abstract: A computer-implemented method. The method comprising: (a) obtaining a candidate characterising code, the candidate characterising code being indicative of character of an example of a physical system; (b) accessing a database of extant characterising codes, the extant characterising codes being indicative of character of other examples of the physical system; (c) determining a degree of similarity between the candidate characterising code, and at least one of the extant characterising code; and (e) providing, as an output, the results of the comparison.

Abstract: The present application provides a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: passing UV-visual spectrum light through the fuel; measuring a transmittance parameter indicating the transmittance of light through the fuel; determining one or more fuel characteristics of the fuel based on the transmittance parameter; and communicating the one or more fuel characteristic to a control module of the gas turbine engine or the aircraft. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: The present application discloses a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method comprises: determining, during use of the gas turbine engine, one or more exhaust content parameters by performing a sensor measurement on an exhaust of the gas turbine engine; and determining one or more fuel characteristics of the fuel based on the one or more exhaust parameters including the nvPM content of the exhaust. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.

Abstract: A single-piece annular vane array for a gas turbine engine. The single-piece annular vane array comprises at least one mounting feature, a ring of stator vanes, and at least one baffle. The at least one baffle is attached at a first radial location to the ring of stator vanes, and is attached at a second radial location to the at least one mounting feature. The at least one baffle is deformable.

Abstract: A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

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: 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: 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 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: 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: 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: 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 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 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: 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.

Abstract: An annulus filler for mounting to a rotor disc of a gas turbine engine includes a coupling portion and an outer lid coupled to the coupling portion. The outer lid includes a leading edge, a trailing edge, and a main portion including an outer radial surface. The outer lid further includes a protruding portion connected to the main portion and extending radially outwardly from the main portion. The protruding portion is axially disposed between and spaced apart from the leading edge and the trailing edge. The protruding portion includes a protruding surface contiguous with and extending radially outwardly from the outer radial surface, such that the protruding surface is configured to redirect air drawn through the gas turbine engine.

Abstract: A method for scanning at least one component includes placing the at least one component inside a storage vessel; filling the storage vessel with a cryogenic material; cooling the at least one component; placing the storage vessel between an x-ray source and a detector of a CT scanner; generating, via the x-ray source, an x-ray cone beam after cooling of the component that passes through the storage vessel while the at least one component is disposed within the storage vessel; receiving the x-ray cone beam at the detector; and generating an x-ray image of the at least one component.