Abstract: Described herein is a non-contact excitation system for a spin rig comprising: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements comprise magnets; and wherein, in use, the exciter elements are arranged around the test piece; wherein each exciter element is axially and radially offset from the plurality of test piece elements. Also described is a method of testing a test piece using the system.

Abstract: A battery cooling block, including: a housing defining a flow chamber and including a first and second port configured to allow flow of coolant through flow chamber along a first axis; the housing further including a plurality of compartments disposed in the flow chamber, each compartment extending along a respective compartment axis perpendicular to first axis and each compartment configured to receive respective cylindrical battery cell; wherein each compartment is defined by partially open compartment wall within flow chamber, compartment wall including a plurality of wall segments extending parallel to compartment axis and which are circumferentially spaced apart around compartment axis wherein respective battery cell is partially exposed to flow of coolant. A battery assembly, including: plurality of cylindrical battery cells; and battery cooling block as described above.

Abstract: A non-contact excitation system for a spin rig includes: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements include magnets; each exciter element being moveable between an active position, in which a magnetic force is provided between each exciter element and at least one of the test piece elements, and an inactive position, in which the magnetic force between the exciter magnets and the at least one test piece element is reduced.

Abstract: An engine for an aircraft includes 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. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.

Abstract: A method for scanning of an object in a scanning apparatus includes disposing the object and a complementary object on a support of the apparatus, so that the objects are positioned between an imaging beam emitting element and an imaging beam receiving element oppositely disposed to either side of the support. The support is rotatable relative to the emitting and receiving elements about an axis of rotation to allow creation of an image from projections each taken at a different relative angle. A volume of the complementary object is solid or filled with a filling material and the complementary object is configured to reduce the variation in imaging beam attenuation across the objects or the part to be scanned of the object at the multiple relative angles of rotation. The method includes operating the scanning apparatus at the multiple relative angles of rotation to produce an image of the object.

Abstract: There is provided an air pressurisation system for an aircraft. The air pressurisation system comprises a blower compressor, a delivery line and a catalyst material. The blower compressor is configured to be mechanically coupled to a spool of a gas turbine engine and configured to receive an inlet flow of gases from a bypass duct of the gas turbine engine. The delivery line is configured to convey gases received from the blower compressor to an airframe port for supply to an airframe system. The catalyst material is disposed along the delivery line and configured to catalyse a reaction of volatile organic compounds within gases conveyed by the delivery line.

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 fuel management system for a gas turbine engine, the fuel management system comprising a fuel supply line configured to supply fuel from a system inlet to a combustor of the gas turbine engine via a combustor pump and combustor valve. A heat exchanger is configured to reject heat from a thermal load of the gas turbine engine to fuel in the fuel supply line between the system inlet and the combustor valve. There is a downstream recirculation line configured to recirculate a downstream excess portion of fuel from the fuel supply line, the downstream recirculation line extending from a downstream recirculation point on the fuel supply line between the heat exchanger and the combustor valve.

Abstract: A ceramic matrix composite (CMC) component is provided that includes: a CMC body in which an environmental protection layer is completely embedded within a CMC material of the CMC body, the environmental protection layer comprising a ceramic that has a higher impact and/or environmental resistance than the CMC material. Methods for manufacturing the CMC component are also provided.

Abstract: A fuel characteristic determination system and method for determining one or more fuel characteristics of an aviation fuel for powering a gas turbine engine of an aircraft. The system includes a sensor component that is formed from a nitrile seal material and includes a surface that can be exposed to the fuel; a sensor for measuring a swell parameter of the seal material; and a fuel characteristics determination module for determining one or more fuel characteristics of the fuel based on the swell parameter. The method includes: exposing one or more seals of a fuel system of an aircraft to fuel within the fuel system; exposing a sensor component, made from the same material as the one or more seals, to the fuel; and measuring a swell parameter of the seal material.

Abstract: A method of operating a gas turbine engine having an engine core having 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 gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a heat exchange system having at least one fuel-oil heat exchanger arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located downstream of the at least one fuel-oil heat exchanger. The method comprises 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: The present disclosure provides a podded engine adaptor configured to secure a podded engine to an aircraft, comprising: an airframe mounting interface configured for attachment to an airframe of an aircraft; an engine mounting interface configured for attachment to an engine; wherein the podded engine adaptor is configured to be attached to an engine at the engine mounting interface and to an airframe of the aircraft at the airframe mounting interface, so as to indirectly couple an engine to the aircraft via the podded engine adaptor.

Abstract: A coating system includes a bond layer including CoNiCrAlY; a thermal barrier layer at least partially disposed on the bond layer and including yttria stabilised zirconia; a mesh layer connecting the thermal barrier layer to the bond layer and including CoNiCrAlY; and an abradable layer disposed on the thermal barrier layer and including magnesium spinel. The mesh layer includes a plurality of cells connected to each other and is at least partially embedded in the thermal barrier layer. Each cell defines a cell opening therethrough and the cell opening of each of the plurality of cells at least partially receives the thermal barrier layer therein. The abradable layer and the thermal barrier layer are at least partially removable from the mesh layer and the bond layer without damaging them.

Abstract: A computer-based control system for a gas turbine engine includes: a controller including control logic, the controller obtains a set of inputs formed by measurements of one or more process variables and values of one or more engine operation set points, and the control logic determines one or more process command values gas turbine engine operation in response to the set of inputs; and an event detection unit obtaining further measurements of one or more process variables and determine whether an abnormal event has occurred based on the further measurements of one or more process variables. The controller also updates the tuning variables using an event accommodation data array, a given tuning variable being changed in response to detection of a given abnormal event when the respective element in the event accommodation data array is assigned an intervention value.

Abstract: A method of operating a gas turbine engine including an engine core and 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, the main gearbox including gears and journal bearings; a recirculating lubrication system arranged to supply oil to lubricate the gears and journal bearings, the lubrication system including a first oil tank arranged to supply oil to the gears and journal bearings and a second oil tank arranged to supply oil to the journal bearings only. The method includes, at cruise conditions: transferring 200-600 kJ/m3 of heat to the fuel from the oil through the heat exchange system so as to control the oil temperature.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A battery pack includes: a plurality of cylindrical battery cells, each one of the plurality of cells having a cylindrical outer cell wall that extends in a z-direction between opposed first and second ends of the cell; a carrier structure that locates the battery cells relative to one another and arranges them in an array in an x-y plane perpendicular to the z-direction; and a plurality of barriers for preventing propagation of thermal runaway between cells of the array, each one of the plurality of barriers forming a barrier wall that surrounds a group of one or more of the cells. A gap is defined between the barrier wall of each respective barrier and the outer cell wall of each of the respective one or more cells enclosed by the respective barrier.

Abstract: An estimation of junction temperatures of transistors in power electronics converter.

Abstract: The disclosure relates to estimation of junction temperatures of transistors in a power electronics converter. Example embodiments include a method of estimating a junction temperature of a transistor in a power electronics converter configured to convert between first and second supply voltages, the method comprising: providing a gate switching signal to the transistor; measuring a rate of change of current through the converter during a switching period of the transistor; and outputting an estimated junction temperature of the transistor based on the measured rate of change of current, wherein the rate of change of current is measured while a gate voltage of the transistor is above a gate threshold voltage and a drain-source voltage across the transistor is above a predetermined fraction of the first or second supply voltage whereby the rate of change of current is measured in a linear region.

Abstract: A method of thrust control for an airbreathing jet engine includes obtaining a demanded thrust setting, a value of static pressure at a first axial location of a combustion chamber of the airbreathing jet engine and a value of static pressure at a second axial location of the combustion chamber of the airbreathing jet engine. The second axial location is downstream of the first axial location. The method also includes obtaining a ratio of the value of static pressure at the first axial location of the combustion chamber to the value of static pressure at the second axial location of the combustion chamber; and controlling a fuel flow rate of the airbreathing engine based at least in part on the demanded thrust setting and the ratio of the value of the static pressure at the first axial location to the value of the static pressure at the second axial location.