Abstract: A computer-implemented method comprising: receiving data comprising two-dimensional data and three-dimensional data of a component of an engine; identifying a feature of the component using the two-dimensional data; determining coordinates of the feature in the two-dimensional data; determining coordinates of the feature in the three-dimensional data using: the determined coordinates of the feature in the two-dimensional data; and a pre-determined transformation between coordinates in two-dimensional data and coordinates in three-dimensional data; and measuring a parameter of the feature of the component using the determined coordinates of the feature in the three-dimensional data.

Abstract: A method comprising: inspecting an engine during a first period of time to identify damage, the engine being associated with an aircraft; receiving three-dimensional data of one or more components of the engine, the three-dimensional data being generated during the first period of time; determining, during the first period of time, whether the identified damage exceeds a threshold; providing instructions to release the aircraft for operation in a second period of time, subsequent to the first period of time, if the identified damage does not exceed the threshold; and inspecting the received three-dimensional data during the second period of time to measure damage.

Abstract: A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged to have the plurality of planet gears (32) mounted thereon; and a gearbox support (40) arranged to at least partially support the gearbox within the engine. The gearbox (30) has a cross sectional area, the cross sectional area being greater than or equal to 2.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Abstract: Method of manufacturing an integral thermoset infused fibre reinforced composite, structural stator vane ring for a core inlet, a bypass duct, or an air intake of a gas turbine engine. The method comprises winding fibre reinforcement material around a mandrel to form an inner annulus preform; providing a plurality of vane preforms comprising fibre reinforcement material, arranging the plurality of vane preforms around the inner annulus preform, and connecting each of the plurality of vane preforms to the inner annulus preform using a fibre jointing method; winding fibre reinforcement material around the plurality of vane preforms to form an outer annulus preform and connecting the outer annulus preform to each of the plurality of vane preforms using a fibre jointing method to produce a stator vane ring preform; and infusing a thermoset resin into the stator vane ring preform and curing the resin to form the integral stator vane ring.

Abstract: Tool arrangement for compacting a composite preform assembly comprising a support structure preform and an array of component preforms each extending from the support structure preform and spaced apart along the support structure preform. The tool arrangement comprises a support tool defining a lay-up surface for laying up the support structure preform; a plurality of component moulds, each component mould comprising a pair of blocks configured to cooperate with one another to receive and compact a respective component preform extending from the support structure therebetween, each block having a compaction surface for engaging the component preform and a driving surface.

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the ? ? maximum ? ? take - off ? ? rotational ? ? speed ? ? of ? ? the ? ? fan fan - turbine ? ? radius ? ? difference ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (?scarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (?droop) relative to the longitudinal centre line. A relationship between the droop angle (?droop) and the scarf angle (?scarf) is given by: ?droop=?scarf/1.5±1 degree.

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, and arrangements of turbine blades around an exhaust nozzle of an engine core. Example embodiments include a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the engine core comprising an inlet upstream of the compressor and an exhaust nozzle at a downstream outlet of the turbine; a fan located upstream of the engine core inlet; and a set of exhaust nozzle vanes spanning the exhaust nozzle, the turbine comprising a first row of turbine blades upstream of the exhaust nozzle vanes and a second row of turbine blades downstream of the exhaust nozzle guide vanes, one or more of the exhaust nozzle guide vanes comprising a passage configured to direct airflow downstream from the first row of turbine blades towards the second row of turbine blades.

Abstract: An engine for an aircraft has 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 for an aircraft is arranged to receive an input from a core shaft and to output drive to a fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and has a sun gear, a plurality of planet gears, a ring gear, and a planet carrier having a plurality of pins, each pin being arranged to have a planet gear of the plurality of planet gears mounted thereon. A ratio of planet carrier torsional stiffness to pin stiffness is within a specified range.

Abstract: A rotor for an electrical machine is provided. The rotor comprises: a rotor body; one or more magnets arranged around the rotor body; and a non-magnetic containment sleeve positioned radially outwardly of the one or more magnets. The containment sleeve has axially-alternating solid ring sections and reticulated ring sections.

Abstract: A gas turbine engine including: an engine core including turbine, compressor, and core shaft connecting the turbine to the compressor; a fan; an epicyclic gearbox that: (i) receives input from the core shaft to drive the fan at a lower rotational speed than the core shaft, and (ii) includes a sun gear, ring gear, planet carrier, and planet gears; and a gearbox support. The gearbox has a cross sectional area greater than or equal to 2.4×10?1 m2. A first gearbox support strength ratio of: torsional ? ? strength ? ? of ? ? gearbox ? ? support radial ? ? bending ? ? stiffness ? ? of ? ? gearbox ? ? support ? × cross ? ? sectional ? ? area ? ? of ? ? gearbox is greater than or equal to 7.

Abstract: A vertical take-off and landing aircraft is shown. The VTOL aircraft has a main wing having a left wing and a right wing configured as folding wings, and one or more of a foreplane, having a left canard and a right canard configured as folding wings, and/or a tailplane, having a left stabiliser and a right stabiliser configured as folding wings. Each one of the folding wings has a fixed inboard section and a folding outboard section. The folding outboard section is downwardly foldable to a landing condition to support the aircraft on a surface.

Abstract: A method of determining the displacement of a component within a device during operation of the device, the method comprising the steps of: obtaining a first x-ray image of the device while the device is in a first operation state; obtaining a second x-ray image of the device while the device is in a second operation state different to the first operation state; processing each of the first and the second image, wherein the processing comprises applying a filter obtained based on the noise of the image and a frequency characteristic of the image; superimposing the first and the second images to align a predetermined point in each of the first and the second images; and measuring the displacement of an edge associated with the component between the first and the second image to obtain the displacement of the component within the device during operation of the device.

Abstract: A fan stage of a ducted fan gas turbine engine has a rotor hub having a principal axis of rotation and a plurality of fan blades having a hub end attached to the hub and extending radially towards a tip end so as to define a blade span dimension. Each blade has a leading and a trailing edge, a chord for a section of the blade being a straight line joining the leading and trailing edges within the section. A difference between a stagger angle in a mid-span region and in the vicinity of the tip end of each blade is greater than or equal to 20°. The fan blades are twisted to a greater extent than conventional between the mid-span and tip end. A camber angle difference between the mid-span region and the tip end may be greater than 30 degrees.

Abstract: A carbon capture system comprising a gas turbine with a heat exchanger operable to heat a working fluid in the gas turbine, a source of high temperature exhaust gas operable to supply heat to the gas turbine through heat exchanger to heat the working fluid wherein the source of high temperature exhaust gas is operable to provide exhaust gas at a high pressure which is greater than the vapor to liquid transition pressure of CO2 at the temperature of a coolant.

Abstract: A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device, an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. The input shaft device having a high rigidity or the input shaft device having a means for decreasing the rigidity, in particular a diaphragm section.

Abstract: An abradable sealing element comprises a substrate and a sealing structure. The sealing structure comprises one or more wall structures extending from the substrate and defining at least one open cell which is filled with abradable material. The one or more wall structures are formed by additive-layer, powder-fed, laser-weld deposition onto the substrate. The one or more wall structures are formed from nickel-based superalloy and constitute from about 10% to about 50% of the total volume of the sealing structure.

Abstract: The present disclosure concerns removal of entrained contaminant particles in a coolant airflow for a gas turbine engine. Example embodiments include a coolant airflow assembly for a gas turbine engine, comprising: a coolant feed passage connected between a supply of coolant air and an inlet of a component to be cooled, the coolant feed passage defining a coolant airflow path and comprising first and second opposing internal faces (305, 306), the inlet of the component connected to the coolant airflow path through one of the first and second internal faces (305, 306) of the coolant feed passage; and a particulate filter for removing entrained particles from the coolant airflow path, comprising: a first filter panel extending from the first face into the coolant airflow path upstream of the inlet of the component; and a second filter panel extending from the second face into the coolant airflow path upstream of the first filter panel.