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 comprising: determining a mass of a fuel being loaded, or which has been loaded, onto the aircraft; determining a corresponding volume of the fuel; determining one or more fuel characteristics of the fuel 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 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 planetary gearbox device includes a static ring gear, a rotatable planet carrier, a plurality of planet gears, each planet gear being connected to a bearing device, wherein at least one of the bearing devices is coupled to at least one lubricant scooping device and/or at least one lubricant reservoir device for collecting lubricant, in particular oil coming from the at least one bearing device, and at least one lubricant channel allows a flow of the collected lubricant towards a lubrication location and/or a lubricant supply for the planetary gearbox device.

Abstract: Disclosed is a fuel system for a gas turbine engine. The system comprises a fuel pump for fluid communication with a fuel reservoir; a driving turbine for driving the fuel pump; and a source of compressed air flow to drive the driving turbine. The source of compressed air may be the engine core, a dedicated fuel system compressor or the compressor of a cabin blower system.

Abstract: A method for manufacturing a composite component includes providing a tool switchable between a deposition configuration and a flange-forming configuration. The tool includes a first portion including a first deposition surface, a second portion including a second deposition surface and a pre-form support surface, and a flange-forming block including a curved surface. The method further includes depositing a composite material at least partially on each of the first deposition surface and the second deposition surface in the deposition configuration to provide a composite pre-form. The method further includes moving the second portion relative to the first portion to transition the tool to the flange-forming configuration, which causes the composite pre-form to at least partially engage each of the pre-form support surface and the curved surface. The curved surface defines a radius of a flange of the composite component.

Abstract: The present disclosure relates to an airframe assembly comprising an airframe support structure defining an engine void and configured to support a gas turbine engine along an engine axis of the engine void; a thermal enclosure configured to surround the gas turbine within the engine void; a heat exchange channel formed between an outer flow-bounding wall of the engine void and an inner wall defined by the thermal enclosure, configured to convey a cooling flow through the engine void outside of the thermal enclosure for the gas turbine engine; and a heat exchanger mounted to the airframe support structure for heat exchange between the cooling flow and a working fluid provided to the heat exchanger.

Abstract: The present disclosure relates to a combustor assembly for a gas turbine engine. The combustor assembly comprises a combustor liner, a combustor head, a cowl and a fastener. The combustor liner and the cowl define a cavity. The combustor liner has an integral lug that extends into the cavity from a wall of the combustor liner. The fastener extends into the combustor liner lug to fasten the combustor head to the combustor liner.

Abstract: There is provided an electrical power system comprising: a DC voltage source, a DC electrical network, a DC to AC to DC converter having a primary side connected to the DC voltage source and a secondary side connected to the DC electrical network, and a controller configured to control the DC to AC to DC converter, wherein the controller is configured to: monitor an electrical current or voltage between the DC voltage source and the DC electrical network; determine, based on the monitored electrical current or voltage, whether the DC electrical network is in a fault condition; and increase a switching frequency of the primary side of the DC to AC to DC converter in response to a positive determination that the DC electrical network is in a fault condition.

Abstract: A method of monitoring turbine blade creep in a gas turbine engine is provided. The method includes: receiving stereo images of a turbine blade of a row of turbine blades, the images having been obtained using a stereo borescope located in the engine adjacent the row of turbine blades; identifying same features of the blade in each of the stereo images; mapping each of the identified features by triangulation onto a 3D space to produce a 3D depth map of at least part of the blade; providing a 3D reference model of the blade; and comparing the 3D reference model with the 3D depth map to measure one or more deviations in shape of the blade to determine an amount of creep-induced distortion of the blade.

Abstract: A combined gas turbine engine and hydrogen fuel cell system includes a hydrogen fuelled gas turbine engine, a cryogenic liquid hydrogen fuel tank, a first fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, a second fuel offtake arranged to divert a portion of hydrogen fuel from the main fuel conduit downstream of the heat exchanger, and a hydrogen fuel cell configured and arranged to produce electric power using hydrogen fuel diverted from the second fuel offtake.

Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting the turbine and compressor; a fan upstream of the engine core including a plurality of fan blades; and a gearbox that receives an input from a gearbox input shaft portion of the core shaft and outputs drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon, and wherein the sun gear receives input from the core shaft. At cruise conditions the product of the torque on the core shaft and radial bending stiffness or tilt stiffness of the gearbox input shaft portion is within a specified range.

Abstract: An electrical system for connecting a rotary electric machine to de networks operating at different voltages V and W where V>W, the electric machine has N=4 polyphase winding sets. The electrical system includes a first set of N=4 ac-dc converter circuits connected in a modular multilevel configuration, each ac-dc converter circuit having a respective index n=(1, . . . , 4) and an ac interface for connection with a corresponding nth winding set, and in which the modular multilevel configuration has P=5 dc outputs each having a respective index p=(1, . . . , 5), and a set of N=4 dc-dc converter circuits each having a respective index n=(1, . . . , 4) and being configured to convert dc power between a voltage V/4 at a first dc interface and a voltage W at a second dc interface.

Abstract: A trans-critical thermodynamic system includes an expansion device and a separator. The expansion device receives a supercritical fluid containing solutes. The expansion device is operable to expand the supercritical fluid to produce a sub-critical gas by reducing a temperature and/or a pressure of the supercritical fluid. The separator removes the solutes from the sub-critical gas.

Abstract: A butterfly valve for a conduit defining a passage for a flow of a fluid therethrough in a flow direction. The butterfly valve includes a shaft rotatably mounted to the conduit and defining a longitudinal axis along its length. The butterfly valve includes a valve body coupled to the shaft, such that the valve body is rotatable along with the shaft about the longitudinal axis between a closed position and a fully open position. The valve body includes a first major surface, a second major surface opposite to the first major surface, a perimeter surface, a central plane, a first lobe, a second lobe, and a third lobe.

Abstract: A combined cycle power generation and storage system is shown. A liquid air energy storage system uses excess power to liquefy air and stores it in a liquid state. A combustion engine produces power by combustion of a carbon-based fuel along with an exhaust stream containing carbon dioxide. A heat recovery system exchanges heat from the exhaust stream to air from the liquid air storage system, and thereby produce a cooled exhaust stream and heated air. An air expansion machine recovers power by expansion of heated air from the heat recovery system. A separation system separates carbon dioxide from ambient air prior to liquefaction during operation of the liquid air energy storage system, and separates carbon dioxide from the cooled exhaust stream during operation of the combustion engine prior to emission of the cooled exhaust stream to atmosphere.

Abstract: There is provided a method of starting a gas turbine engine, comprising: selecting a baseline starting procedure of the gas turbine engine based on environmental data; receiving condition information for the gas turbine engine and/or for a starting system of the gas turbine engine, the condition information comprising at least one of: health information for the gas turbine engine and/or the starting system; maintenance information for the gas turbine engine and/or the starting system; and operation information relating to an expected future usage of the gas turbine engine. The method further comprises determining a predicted degradation profile for the gas turbine engine and/or the starting system based on the condition information; determining a starting procedure of the gas turbine engine by adapting the baseline starting procedure based on the predicted degradation profile; and controlling the gas turbine engine and/or the starting system according to the determined starting procedure.

Abstract: A gas turbine engine for an aircraft configured with an engine core that has a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, that has a plurality of fan blades. A gearbox arranged to receive an input from the core shaft and to output to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox having a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The gearbox having an overall gear mesh stiffness, and wherein the overall gear mesh stiffness of the gearbox is greater than or equal to 1.05×109 N/m and less than or equal to 8.0×109 N/m.

Abstract: A gas turbine engine for an aircraft includes a staged combustion system having pilot fuel injectors and main fuel injectors. The gas turbine engine further includes a fuel delivery regulator arranged to control delivery of fuel to the pilot and main fuel injectors, and a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being supplied to the staged combustion system. A controller is configured to determine a staging point defining the point at which the staged combustion system is switched between pilot-only operation and pilot-and-main operation, the staging point being determined based on the determined one or more fuel characteristics, the controller being configured to control the staged combustion system according to the determined staging point.

Abstract: Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of: the ? radial ? bending ? stiffness ? of ? at ? least ? one ? of the ? fan ? shaft ? at ? the ? output ? of ? the ? gearbox and ? the ? gearbox ? support the ? moment ? of ? inertia ? of ? the ? fan may be greater than or equal to 2.5×10?2 Nkg?1 m?1 mm?2.

Abstract: There is disclosed an energy storage system for an electric aircraft, the energy storage system comprising: at least one battery pack configured to be disposed onboard the aircraft, and a thermal management system. The thermal management system comprises a first circulation loop configured to be disposed onboard the aircraft and configured to contain a first working fluid, the first circulation loop including: a variable speed pump configured to pump the first working fluid around the first circulation loop, a battery heat exchanger configured to provide a thermal interface between the at least one battery pack and the first working fluid, and a controller configured to control operation of the variable speed pump to intermittently pump the first working fluid to distribute heat around the thermal management system.