Abstract: A method of monitoring a health state of a power semiconductor device includes: receiving an initial transient voltage parameter relating to transient voltage of the power semiconductor device when turning on, when it is known to be unaged; receiving an initial transient current parameter relating to transient current of the power semiconductor device when turning on, when it is known to be unaged; determining an initial energy parameter, relating to initial turn-on switching energy or power of the power semiconductor device, when it is known to be unaged; receiving an operating transient voltage parameter relating to the transient voltage of the power semiconductor device in operation; receiving an operating transient current parameter relating to the transient current of the power semiconductor device in operation; determining an operating energy parameter; and determining the health state of the power semiconductor device based on the initial energy parameter and the operating energy parameter.

Abstract: A variable-geometry convergent-divergent exhaust nozzle includes: a duct to receive exhaust from a combustor; and first and second flap assemblies each including proximal and distal flaps. The first and second proximal flaps are rotatably coupled to the exhaust duct and include entry surfaces that in part define a nozzle passageway to convey the exhaust to an exterior of the engine. The first and second distal flaps are rotatably coupled to the first and second proximal flaps. The first distal flap includes converging and diverging surfaces angled with respect to each other. The second distal flap includes converging and diverging surfaces angled with respect to each other. The converging surfaces define at least in part a convergent portion of the nozzle passageway. The first and second diverging surfaces define at least in part a divergent portion of the nozzle passageway.

Abstract: A method of operating a gas turbine engine having a spool includes executing, by a controller, a shutdown procedure. The shutdown procedure includes: performing an idle rotation operation; determining a parameter associated with a thermal condition of the gas turbine engine while the idle rotation operation is performed; determining, based on the determined parameter, whether the gas turbine engine has met a pre-determined criterion corresponding to a thermally stabilised condition; and terminating the idle rotation operation in response to a determination that the gas turbine engine has met the pre-determined criterion.

Abstract: A propulsion system having an axial direction relating to a principal direction of propulsion, and a height direction perpendicular to the axial direction. The propulsion system includes a duct configured to convey a flow of fluid along the axial direction, and a heat exchanger disposed within the duct. The heat exchanger is defined between an inlet face and an outlet face and is configured to convey the flow of fluid from the inlet face to the outlet face. A heat exchanger length is defined as a minimum distance between the inlet face and the outlet face and wherein the heat exchanger length varies along the height direction such that a resistance to the flow of fluid through the heat exchanger varies along the height direction.

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 torque on the core shaft is greater than 10,000 Nm and a ratio of core shaft stiffness to core shaft torque is within a specified range.

Abstract: A section of a continuum arm robot including a pair of pivot points at a first and second end of the joint, the section further having a first and second connecting rod connected to the section and joining at the centre of the section, the connecting rods has a plurality of gear teeth at the peripheral ends of the connecting rod that intermesh at a central point on the section.

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 method of operating a gas turbine engine having a spool includes executing, by a controller, a pre-shutdown procedure. The pre-shutdown procedure includes: determining a parameter associated with a thermal condition of the gas turbine engine; and calculating an idle period for an idle rotation operation to be performed in a shutdown procedure. Calculation of the idle period is based on the determined parameter.

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 complex cycle gas turbine engine for an aircraft with hydrogen fuel supply. The gas turbine engine comprises, in fluid flow series, a core gas turbine and a recuperator for heating hydrogen fuel prior to combustion by turbine exhaust products.

Abstract: A cooling system for an aircraft comprises a gas turbine engine, an ancillary apparatus, and a heat exchanger. The gas turbine engine comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate an electrical power PEM1 (W). The heat exchanger is configured to transfer a total waste heat energy Q (W) generated by the gas turbine engine and the ancillary apparatus, to an airflow passing through the heat exchanger, and a ratio S of: S = ( Total ? Electrical ? Power ? Generated = P EM ? 1 ) ( Total ? Heat ? Energy ? Rejected ? to ? Airflow = Q ) is in a range of between 0.50 and 5.00.

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. 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 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 includes operating the gas turbine engine using a fuel having a lubricity of between 0.71 mm and 0.90 mm wear scar diameter (WSD) at 25° C.

Abstract: A method of producing a freestanding ceramic tile. The method involves: grit-blasting a substrate using a grit size in the range from 36 to 220 mesh; depositing a release layer of carbon or graphite from 2 to 10 microns thick on the grit-blasted surface of the substrate; applying ceramic over the release layer until a desired thickness of ceramic is achieved to form a ceramic layer; heating the substrate, release layer, and ceramic layer to a temperature of from 800 to 1000 degrees Celsius; keeping the substrate, release layer, and ceramic layer at a temperature from 800 to 1000 degrees Celsius for a time from 10 to 20 minutes to remove the release layer; and cooling the substrate and ceramic layer at a rate of at least 200 degrees Celsius per minute, such that the ceramic layer separates from the substrate to produce a freestanding ceramic tile.

Abstract: Disclosed is a gas turbine engine rotor balancing apparatus comprising: a rotor configured to rotate about a rotor axis; a substantially annular balancing mass configured to encircle the rotor; and a mass adjustment apparatus for adjusting the position of the balancing mass relative to the rotor.

Abstract: The disclosure relates to thermal management of a power electronics converter. An example embodiment includes a power electronics converter comprising: a first set of terminals connectable to a first voltage supply; a second set of terminals connectable to a second voltage supply; a plurality of semiconductor switches connected between the first and second sets of terminals; a voltage sensor connected to measure a voltage across the second supply; a current sensor connected to measure a current through the second supply; and a controller connected to provide switching signals to the semiconductor switches and to receive voltage and current signals from the voltage and current sensors, the controller configured, for each of the semiconductor switches, to: determine a junction temperature of the semiconductor switch; compare the determined junction temperature to a reference temperature; and adjust a switching gate voltage applied to the semiconductor switch.

Abstract: There is disclosed a method of manufacturing a composite component comprising a main body and an integral flange, the method comprising applying fibre-reinforcement material on a tool having a main body portion and a flange-forming portion to provide a pre-form comprising a body region and a longitudinally adjacent flange region. The pre-form extends generally longitudinally between two longitudinal ends; and a trailing ply of the pre-form extends generally longitudinally between the longitudinal end closest to the flange region and an inner ply end located in the flange region or partway into the body region. Relative movement between the flange-forming portion and the main body portion causes sliding movement between the trailing ply and the flange-forming portion during a flange forming operation, thereby causing a tension force in at least the flange region of the pre-form of during forming of the flange.

Abstract: A continuum arm robot system comprising at least a first continuum arm robot and a second continuum arm robot, each continuum arm robot being controlled by its own actuator pack, and each actuator pack being coupled to a single control computer, wherein at least the second continuum arm robot comprises a releasable connection mechanism to engage in gripping the first continuum arm robot in a workspace, so as to link the at least two continuum arm robots into a single redundant robotic system with at least the second continuum arm robot providing support for the first continuum arm robot.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: An exhaust nozzle for a gas turbine engine includes: an exhaust duct configured to receive an exhaust flow of gas from a combustor of the engine; a first flap rotatably coupled to the exhaust duct for rotation about a first axis; a first actuator configured to actuate the first flap about the first axis between a first inner and a first outer position; a second flap rotatably coupled to the exhaust duct for rotation about a second axis; and a second actuator configured to actuate the second flap about the second axis between a second inner and a second outer position. The first and second flaps at least in part define a passageway configured to convey the exhaust flow of gas to an exterior of the gas turbine engine. The first and second axes of rotation are coaxial. Also provides is a method of operating an exhaust nozzle.