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

Abstract: A gas turbine engine configured with an engine core. A fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear. Additionally, the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×1018 N2m?2.

Abstract: An article and a method of manufacturing the article is disclosed. The method includes providing the article including a substrate and a coating at least partially disposed on the substrate. The coating includes an outer surface. The coating further includes platinum and chromium. The method further includes applying cold work to the outer surface of the coating to produce a cold-worked layer extending from the outer surface of the coating to a cold work depth. The cold-worked layer includes approximately 45% cold work. The cold work depth is between about 30 microns to about 150 microns from the outer surface of the coating.

Abstract: An aircraft comprises a hydrogen-fuelled propulsion system, a plurality of like generally cylindrical hydrogen storage tanks and a conveying system arranged to convey hydrogen from the hydrogen storage tanks to the hydrogen-fuelled propulsion system. The aircraft further comprises a fuselage having a cargo bay (502) including one or more (510A-G) of the plurality of hydrogen storage tanks, the longitudinal axes (511A-G) of the one or more hydrogen storage tanks within the cargo bay extending parallel to the longitudinal axis (501) of the fuselage and lying in one or more planes (595, 597) extending across the width dimension of the cargo bay. The hydrogen storage tanks within the cargo bay have a common aspect ratio R in the range 4.2 ? R ? 25.7, allowing the volume of space with the cargo bay occupied by stored hydrogen to be maximised or approximately maximised.

Abstract: A gas turbine engine includes: a fan rotating about an engine main axis; a core duct; an engine core; an Engine Section Stator (ESS) including a plurality of ESS vanes and arranged in the core duct downstream of the fan; and a plurality of variable inlet guide vanes (VIGV) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS. The VIGV vanes are arranged angularly rotated with respect to the ESS vanes such that the VIGVs are shielded by the ESS, thereby protecting the VIGVs from icing and from ice shedding from the ESS vanes.

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: An aircraft cabin blower system comprises a cabin blower including a compressor configured to provide air to a cabin of the aircraft; a variable drive system configured to drive the compressor and including an electric variator and a summing gearbox; and a main transmission configured, when operating in a blower mode, to receive mechanical power from a gas turbine engine and input mechanical power to the summing gearbox in a forward direction; and configured, when operating in a starter mode, to receive mechanical power from the summing gearbox and input mechanical power to the gas turbine engine. The aircraft cabin blower system further includes a first one-way rotation device adapted to permit free rotation of the main transmission in the forward direction and to prevent rotation of the main transmission in a reverse direction opposite to the forward direction.

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 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 turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly includes fan blades defining a corresponding fan area (AFAN). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and includes radially-extending vanes arranged in a circumferential array with at least one vane including a heat transfer element for heat transfer from a first fluid contained within each element to an airflow passing over a surface of each heat transfer element before entering the fan assembly inlet. Each heat transfer element extends axially along the corresponding vane, with a swept heat transfer element area (AHTE) being the wetted surface area of all heat transfer elements in contact with the airflow. A Fan to Element Area parameter FEA of AHTE/AFAN lies in the range of 47 to 132.

Abstract: A roller bearing arrangement for a gas turbine engine. The roller bearing arrangement includes a fan shaft, and a stub shaft connected to the fan shaft. The roller bearing arrangement further includes a plurality of roller bearing elements positioned between a first axial bearing surface created on a radially outer surface of the stub shaft and a second axial bearing surface of a static structure, the roller bearing arrangement further including a first snubber positioned between the radially outer surface of the fan shaft and a radially inner surface of the stub shaft, the first snubber being spaced apart from the radially inner surface of the stub shaft or the radially outer surface of the fan shaft so as to limit a radial movement range of the stub shaft.

Abstract: The present disclosure relates to gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox comprising: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

Abstract: A turbine blade including an aerofoil and a shroud. The shroud includes a first abutment surface configured to face a second abutment surface of a first circumferentially adjacent turbine blade. The shroud further includes a second abutment surface configured to face a first abutment surface of a second circumferentially adjacent turbine blade. The shroud further includes an inner platform surface extending at least circumferentially between the first abutment surface and the second abutment surface. The shroud further includes a first recessed surface extending at least radially and circumferentially from the first abutment surface to the inner platform surface. The first recessed surface defines a first recessed region configured to receive a flow of a cooling fluid from the first circumferentially adjacent turbine blade.

Abstract: A bladed disc for a rotating machine comprising a central disc that rotates about a central axis, the central disc having a series of blades arranged around its periphery; the blades have dovetail roots which engage with slots on the central disc; the bladed disc being configured so that there is a pre-loading force between the blades and the central disc such that each blade is forced away from the central axis of the bladed disc; and wherein the pre-loading force is equal or greater than 40% of the maximum centrifugal force applied to the blade during a flight cycle.

Abstract: An air-breathing turbojet engine for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet configured to compress inlet air by one or more shocks, a cooler to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor to compress the air further. A precooler cools the compressed inlet air using compressed cooled air from the turbo-compressor. A combustor receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine expands and is driven by combustion products, and a second turbine expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

Abstract: A core duct assembly for a gas turbine engine includes a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.

Abstract: A computer-implemented method comprising: receiving a first signal from a first sensor, the first signal having a first value; receiving a second signal from a second sensor, the second signal having a second value, the first signal and the second signal being measurements of a parameter of a system; receiving a third signal having a third value; determining a first absolute error using the first value, the second value and the third value; and comparing the first absolute error with at least one threshold to determine a signal value.

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 propulsion system for an aircraft comprises a gas turbine engine; a plurality of fuel tanks arranged to contain different fuels to be used to power the gas turbine engine, wherein the fuels have different calorific values; and a fuel manager. The fuel manager is arranged to store information on the fuel contained in each fuel tank and to control fuel input to the gas turbine engine in operation by selection of a specific fuel or fuel combination from one or more of the plurality of fuel tanks based on thrust demand of the gas turbine engine such that a fuel with a lower calorific value is supplied to the gas turbine engine at lower thrust demand.

Abstract: A control system for a compliant robotic system including at least two compliant robots having actuator packs, the control system including: an individual local control system associated with each actuator pack, the local control system providing control signals to the actuator causing movement within the robots, an overall control system controlling the overall motion of robots when proximate within a workspace, the overall control signal providing signals to the actuators associated with the robots, so as to cause linked movement of the continuum arm robots, and wherein each individual control system is provided with a clock synchronized with the other, and wherein the overall control system is provided with a redundancy control system that limits the motion of the compliant robots within certain degrees of freedom, so that the motion of the at least two complaint robots does not conflict when operating under the overall control system.