Abstract: A coolant for a subtractive machine process, the coolant comprising a phase change material which changes from a solid state to a liquid state as a result of frictional heat generated during the subtractive machine process of the machine tool on a component to be machined, and returns to the solid state as the component cools.

Abstract: There is provided a method of controlling a robot within an environment comprising: i) receiving, from a 3D scanner, data relating to at least a portion of the environment for constructing a 3D point cloud representing at least a portion of the environment; ii) comparing the 3D point cloud to a virtual 3D model of the environment and, based upon the comparison, determining a position of the robot; then iii) determining a movement trajectory for the robot based upon the determined position of the robot. Also provided is a control apparatus and a robot control system.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A hub-tip ratio of the fan, defined as the ratio of the diameter of the hub to the diameter of the fan measured at the leading edge of the blades, is from 0.45 to 0.55.

Abstract: A mounting apparatus for mounting a gas turbine engine to an aircraft. The mounting apparatus includes a thrust strut operably connected to the gas turbine engine and the aircraft. The thrust strut defines a thrust strut axis. The mounting apparatus further includes a restraining structure. The restraining structure includes a bracket circumferentially disposed on the thrust strut and at least one elongate member connected to the bracket and the gas turbine engine. The restraining structure radially and circumferentially constrains the thrust strut with respect to the thrust strut axis of the thrust strut while allowing the thrust strut to move in the direction of thrust strut axis.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft includes a nacelle (501) defining a duct (502), and a fan (503) located therewithin. The fan comprises a hub which rotates around a rotational axis (A-A) and a plurality of blades attached thereto. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r?[0, 1]), a leading edge and a trailing edge defining, for each span position, a chord therebetween having a chord length (c), and a blade thickness (t) defined for each span position thereof. For each blade, a ratio of thickness at the 0 percent span position (thub) to chord length is 0.1 or greater.

Abstract: A component for a gas turbine engine, comprising: first and second walls; a coolant channel defined by the space between the first and second walls; and a first rib extending between the first and second walls to the end of the coolant channel in a coolant flow direction, such that the coolant channel is bifurcated in the coolant flow direction.

Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Abstract: A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of: the ? ? lowest ? ? frequency ? ? of ? ? either ? ? mode ? ? Fan ? ? RTW ? ? or ? ? Rotor ? ? RW ? ? at the ? ? MTO ? ? speed the ? ? MTO ? ? speed may be in the range from 15 to 50%.

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: An engine core including a turbine, compressor, and a core shaft connecting the turbine and compressor; a fan located upstream of the engine core including a plurality of fan blades; and a gearbox. The gearbox is arranged to receive an input from the core shaft and to output drive to the fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and includes a sun gear, a plurality of planet gears, ring gear, and planet carrier to the mounted planet gears. The planet carrier has an effective linear torsional stiffness and the gearbox has a gear mesh stiffness between the planet gears and the ring gear. A carrier to ring mesh ratio of: the ? ? effective ? ? linear ? ? torsional ? ? stiffness ? ? of ? ? the ? ? planet ? ? carrier gear ? ? mesh ? ? stiffness ? ? between ? ? the ? ? planet ? ? gears ? ? and ? ? the ? ? ring ? ? gear is greater than or equal to 0.2.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, which includes a primary combustion zone with a primary combustion zone depth and a secondary combustion zone. A ratio of the primary combustion zone depth to the lean burn fuel injector head tip diameter is less than 2.4.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: A fuel delivery system (201) is shown for delivering the hydrogen fuel from a cryogenic storage system to a fuel injection system in a gas turbine engine. The fuel delivery system includes a pump (301), a metering device (302), and a fuel heating system (303,304) for heating the hydrogen fuel to an injection temperature for the fuel injection system.

Abstract: An aircraft gas turbine engine includes a fan system having a reverse traveling wave first flap mode, Fan RTW, and including a fan upstream of the engine core; a fan shaft; and a front engine structure to support the shaft and having a front engine structure nodding mode FSN including two modes at similar, but unequal, natural frequencies in orthogonal directions; and a gearbox. The engine includes a gearbox, and a gearbox output shaft to couple output of the gearbox to the fan shaft. An LP rotor system including the fan system and the gearbox output shaft has a first reverse whirl rotor dynamic mode, Rotor RW.

Abstract: A Lubrication system for an engine component of a gas turbine engine, such as a planetary gearbox. The system comprises a main lubrication system comprising a main tank configured to supply lubricant to the engine component, and a main reservoir configured to collect lubricant scavenged from the engine component after lubrication; and an auxiliary lubrication system comprising an auxiliary reservoir configured to collect lubricant scavenged from the engine component after lubrication and to supply scavenged lubricant to the engine component. Main reservoir comprises an interior and the auxiliary reservoir is located within the interior of the main reservoir. The auxiliary reservoir comprises an upper portion, a lower portion and a lubricant pickup region, the upper portion comprising an opening arranged to permit lubricant to overflow into the main reservoir, and the lower portion of the auxiliary reservoir being shaped to cause lubricant to drain towards the lubricant pickup region.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, which includes a primary combustion zone with a primary combustion zone depth and a secondary combustion zone. A ratio of the primary combustion zone depth to the lean burn fuel injector head tip diameter is less than 2.4.

Abstract: There is disclosed an alignment tool for positioning an impact liner panel on a fan casing. The alignment tool comprises an attachment portion for attaching to the fan casing, a support surface for receiving a shim, and a magnet to magnetically retain a shim on the support surface. There is also disclose a tool kit for manufacturing a fan casing having an alignment tool and a shim. There is also disclosed a method of positioning an impact liner panel on a fan casing. It comprises attaching an alignment tool to the fan casing, magnetically retaining a shim on a support surface of the alignment tool; and positioning the impact liner panel against an abutment surface of the shim.

Abstract: A turbine vane assembly includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The turbine vanes are adapted to interact with gases flowing through a gas path of the gas turbine engine. The outer vane support is arranged radially outward of the turbine vane and configured to receive aerodynamic loads from the turbine vanes. The inner vane support is arranged radially inward from the outer vane support to locate the turbine vane assembly therebetween.

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.