Abstract: A resin composition that is useful in the preparation of protective coatings. The resin composition comprises a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer, and an imidazolium dicyanamide catalyst. The method of preparing a resin blend by mixing a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer and an imidazolium dicyanamide catalyst to form a resin composition and curing the resin composition to form the resin blend.

Abstract: The disclosure relates to a heat exchanger for a cooling system, the heat exchanger comprising: a first part having a first surface for connecting to a component to be cooled and a second opposing surface, the first part having a fluid flow channel extending from a fluid inlet through the first part between the first and second surfaces; a second part extending from the second surface to an external third surface, the second part having an open-cell porous structure in fluid communication with the fluid flow channel such that fluid flowing through the fluid flow channel passes into the second part and exits the heat exchanger at the external third surface.

Abstract: The disclosure relates to power electronics converters having circuit breaker protection for use in case of faults. Example embodiments include a power electronics converter for converting an input AC supply having a plurality of phases to an output DC supply, the converter comprising: a plurality of AC input terminals connectable to the plurality of phases of the AC supply; first and second DC output terminals; a capacitor connected between the first and second DC output terminals; a first MOSFET connected between each of the plurality of AC input terminals and the first DC output terminal; a second MOSFET connected between each of the plurality of AC input terminals and the second DC output terminal; and a third MOSFET reverse connected in series with the first MOSFET between the first DC output terminal and each of the plurality of AC input terminals.

Abstract: A combustor assembly for a gas turbine engine includes a combustor shell extending along a shell axis. The combustor shell includes a shell wall extending around the shell axis and a meter panel connected to an upstream end of the shell wall. The shell wall and the meter panel together define a combustion space therebetween. The combustor assembly further includes: a tile disposed within the combustor shell and including a tile flange portion extending radially with respect to the shell axis; a heatshield connected to the meter panel and extending at least radially towards the tile; a fastener connecting the tile flange portion to the meter panel; and a compartment thermally shielded from the combustion space. Either the heatshield and the tile together form the compartment or the tile alone forms the compartment. The fastener is spaced apart from the combustion space and at least partially disposed within the compartment.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? exhaust ? nozzle ? pressure ? ratio core ? exhaust ? nozzle ? pressure ? ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

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 gas turbine engine having a combustor supplied with fuel from a fuel system, the method comprising: determining a mass of the fuel being supplied to the combustor; determining a corresponding volume of the fuel being supplied to the combustor; and determining one or more fuel characteristics based on the determined mass and volume. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A battery pack core may include a cold plate comprising a plurality of apertures defined between a first major surface and a second major surface of the cold plate; a plurality of battery cells, a single battery cell positioned in each aperture of the plurality of apertures such that a first end of the battery cell projects beyond the first major surface and a second end of the battery cell projects beyond the second major surface; and a plurality of silicone bushings, a silicone bushing surrounding each battery cell of the plurality of battery cells and contacting a wall of the aperture in which the battery cell is positioned.

Abstract: An aircraft propulsion system comprises a gas turbine engine arranged to provide propulsive thrust and a fuel cell system having an air input port, the aircraft propulsion system being configured such that air from a compressor of the gas turbine engine is provided to the air input port during operation of the aircraft propulsion system. The fuel cell system is able to provide appreciable electrical power at altitude without the need for a dedicated compressor.

Abstract: A computer-implemented method of enabling optimisation of derate for a propulsion system of a vehicle, the method comprising: determining a derate for the propulsion system of the vehicle using: an algorithm; a vehicle model defining path constraints for the vehicle through space; a propulsion system model defining parameters of the propulsion system; an objective function defining one or more objectives; and controlling output of the determined derate.

Abstract: A method of manufacturing a component includes forming a mould assembly including an initial mould unit, providing a seed crystal including a primary growth direction, determining an optimal angular orientation of the unit, rotating the unit to dispose the unit's optimal angular orientation, encasing the unit in a refractory material, and forming a refractory mould unit having a component mould including a mould wall defining a mould cavity, and a seed holder. In the optimal angular orientation, the seed crystal's primary growth direction is angled away from the wall, thereby forming a converging disposition with the wall in a of the wall's first region facing the central sprue and a diverging disposition with the wall in the wall's second region facing a mould heater. The method includes receiving the seed crystal within the seed holder and filling the mould cavity with molten castable material to form the component.

Abstract: A method of operating a gas turbine engine and a gas turbine engine includes a fuel delivery system arranged to provide fuel, a combustor arranged to combust at least a proportion of the fuel, a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough, and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. A fuel viscosity is adjusted to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: A method of checking refuelling of an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises: receiving an input of calorific value data for fuel provided to the aircraft on refuelling; independently determining at least one of: (i) the calorific value of fuel supplied to the gas turbine engine in use; and (ii) the calorific value of the fuel provided to the aircraft on refuelling; and providing an alert if the determined calorific value is inconsistent with the calorific value data input received.

Abstract: A method of operating a gas turbine engine including: a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor, wherein the fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, which are arranged to transfer heat between the oil and fuel and include primary and secondary fuel-oil heat exchangers; a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the heat exchangers; and a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel which has passed through the primary heat exchanger to be returned to the inlet. The method includes selecting one or more fuels such that the calorific value of the fuel provided to the gas turbine engine is at least 43.5 MJ/kg.

Abstract: A kinetic energy absorptive composite article comprising a plurality of plies, and a method of manufacturing the kinetic energy absorptive composite article. Each ply comprises a plurality of substantially parallel fibers encapsulated within a resin. A plurality of localised weaknesses are comprised within the fibers of the kinetic energy absorptive composite article, the locations of the plurality of localised weaknesses forming a pre-defined pattern. The pre-defined pattern comprises a first sub-pattern and a second sub-pattern superposed upon the first sub-pattern. The first sub-pattern is configured so that a crack formed by an impact event propagates substantially along a predetermined fracture path to separate the kinetic energy absorptive composite article into at least a first portion and a second portion and the second sub-pattern is configured to set a threshold below which the crack does not propagate.

Abstract: An integral guide vane combustor assembly for a gas turbine engine comprises an annular array of radially extending guide vane combustors. The gas turbine engine comprises, in axial flow sequence, a compressor assembly, the integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly. Each of the guide vane combustors comprises, in axial flow sequence, a guide vane portion, and a combustor body portion. The guide vane portion is formed integrally with the combustor body portion, and the guide vane portion is configured to direct a gas flow exiting the compressor assembly into the combustor body portion.

Abstract: A fuel swirl nozzle for a gas turbine engine comprises, in axial fuel flow sequence, a fuel inlet portion, a fuel stem portion, and a fuel/air swirler portion. The fuel inlet portion is configured to receive a flow of a fuel, with an inlet to the fuel stem portion being fluidly sealed against the fuel inlet portion, and an outlet from the fuel stem portion being fluidly sealed against the fuel/air swirler portion. The fuel/air swirler portion comprises an outer air swirler and an inner air swirler, with the inner air swirler being positioned concentrically within the outer air swirler. The outer air swirler comprises a plurality of first vanes that are arranged in an annular array within an outer air swirler body and is configured to impose a first rotational flow component on an air flow entering the outer air swirler.

Abstract: A combustor assembly for a gas turbine engine comprises a combustor casing. The combustor casing has a diametral height D and an axial length L. An Aspect Ratio parameter S is defined as: S = D L and the S parameter lies in the range of 0.30 to 1.00.

Abstract: A fuel swirl nozzle for a gas turbine engine comprises, in axial fuel flow sequence, a fuel inlet portion, a fuel stem portion, and a fuel/air swirler portion. An inlet to the fuel stem portion is fluidly sealed against the fuel inlet portion, and an outlet from the fuel stem portion is fluidly sealed against the fuel/air swirler portion. The fuel/air swirler portion has a cross-sectional profile that is generally dolioform shaped along an axis parallel to the engine's longitudinal axis, and has a first end and a second end. The first end of the fuel/air swirler portion has an air inlet face, and the second end of the fuel/air swirler portion has a fuel/air mixture outlet face. The first end comprises a rim portion enclosing the air inlet face in which the rim portion has a rim portion radius (rrim) normal to an outer circumference of the inlet face.

Abstract: A combustor assembly for a gas turbine engine comprises a combustor casing, and an integer quantity N of fuel swirl nozzles. The combustor casing defines a total internal volume of the combustor casing V (cm3) together with a Fuel Swirl Nozzle Density ratio (m?3) that is defined as: D F ? S ? N = N V and the DFSN parameter is a value in the range of 200 to 1,500.

Abstract: A method of determining one or more fuel characteristic of an aviation fuel suitable for powering a gas turbine engine of an aircraft is disclosed. The method includes: determining, during use of the gas turbine engine, one or more contrail parameters related to contrail formation by the gas turbine engine. The determining of the one or more contrail parameters includes performing a sensor measurement on a region behind the gas turbine engine in which a contrail is or can be formed. The method also includes determining one or more fuel characteristics of the fuel based on the one or more contrail parameters. A fuel characteristic determination system, a method of operating an aircraft and an aircraft are also disclosed.