Abstract: A carrier segment for forming a carrier section for one or more gas turbine engine components, the carrier segment including a carrier wall extending between front and rear carrier ends and having a circumferential profile, the carrier segment being configured for defining radially inwardly thereof a chamber for receiving cooling air from an outboard feed source via the front end, wherein the circumferential profile of the carrier wall is undulating, e.g. with a simple waveform configuration or a more complex arrangement of interleaved series of oppositely oriented part-conical surface formations.

Abstract: A method of providing a threaded fastener with a multi-layered coating comprises the step of using a vapor deposition process to apply a multi-layered coating to a threaded portion of the threaded fastener, each layer of the multi-layered coating being formed from a coating material selected from the group comprising titanium, chromium, nickel, platinum, tantalum, and combinations thereof.

Abstract: There is disclosed a method for the production of a three-dimensional product (2) via an additive layer manufacturing process such as an electron beam manufacturing process to selectively fuse parts (17) of a powder bed (16), said parts (17) corresponding to successive cross-sections of the product (2). The method involves the use of said additive layer manufacturing process to form a tool (12) by selectively fusing additional parts (18) of the powder bed (16), said additional parts (18) corresponding to successive cross-sections of the tool (12). The method also comprises a subsequent step of manipulating the tool (12) perform a processing function on the product (2).

Abstract: Air-riding seal arrangement between two components are rotatable relative to each other about a rotation axis. The seal arrangement includes an axially slidable carrier module mounted to one component and runner provided by other components. The carrier module and runner have respective and opposing annular sealing faces in a seal arrangement normal operation are axially spaced to form an air-seal therebetween separating a high-pressure region at one side of the air-seal from low-pressure region at the other side. The carrier module has a front and rear portions, and carrier module sealing face is provided by the front portion. The front portion has first and second rotation positions relative to the rear portion, the first rotation position being in normal operation, and second rotation position being in event of high speed touchdown of the annular sealing faces when contact between sealing faces rotates the front portion to the second rotation position.

Abstract: An air swirler has a radially outer wall bounding an annular passage. The radially outer wall has an axially extending portion and a radially convergent portion which converges towards an outlet of the passage. Downstream of the outlet, the passage facing surface of the radially outer wall becomes divergent. The region in which the passage facing surface of the radially outer wall is convergent then divergent defines the flare. A circumferential and axially extending recess is provided in the radially outer wall opposite the flare. A reinforcement of material with superior wear properties to the material of the main body is built in the recess to a radius which coincides with that of the outer surface of the axially extending portion of the radially outer wall. The reinforcement can conveniently be added to the main body by means of a laser consolidation ALM process.

Abstract: A shaft has an axis, an outer surface, an inner surface and is arranged to rotate in a first circumferential direction. The shaft comprises a first set of equally circumferentially spaced apertures extending through the shaft between the outer and inner surfaces of the shaft. A second set of equally circumferentially spaced apertures extends through the shaft between the outer and inner surfaces of the shaft. The first and second sets of apertures are arranged in first and second planes arranged perpendicularly to the axis of the shaft. The second plane is spaced axially downstream from the first plane. Each aperture in the second set of apertures is circumferentially displaced in a second circumferential direction from a circumferential mid position between the two apertures in the first set of apertures to reduce the stresses in the shaft.

Abstract: A lean burn fuel injector comprises a fuel feed arm and a fuel injector head. The fuel feed arm has a single pilot fuel supply passage and a single main fuel supply passage extending there-through. The fuel injector head has a coaxial arrangement of an inner pilot air-blast fuel injector and a coaxially arranged outer main air-blast fuel injector. The inner pilot air-blast fuel injector comprises a coaxial arrangement of a pilot inner air swirler passage and a pilot outer air swirler passage. The outer main air-blast fuel injector comprises a coaxial arrangement of a main inner air swirler passage and a main outer air swirler passage. The single main fuel supply passage is arranged to supply main fuel to the main inner air swirler passage. The single pilot fuel supply passage is arranged to supply pilot fuel onto the pre-filming surfaces in the pilot inner and outer air swirler passages.

Abstract: An aerofoil configured to be connected to a second component, the aerofoil comprising a first face including a bi-cast groove for joining the aerofoil to the second component by bi-casting; a second face connected to a main body of the aerofoil; an angled sealing face between the first and second face, the angled sealing face being angled with respect to the second face and configured to provide a sealing interface between the first face and the second face; and a concave face forming a recessed surface at the intersection of the sealing face and the second face.

Abstract: A combustion chamber arrangement comprises an annular combustion chamber, a plurality of circumferentially arranged vanes and a radially inner sealing ring. The vanes are arranged at the axially downstream end of the annular combustion chamber. Each of the vanes comprises an aerofoil extending radially between a radially inner platform and a radially outer platform. The radially inner sealing ring forms a seal between a downstream end of a radially inner annular wall structure of the combustion chamber and the radially inner platforms of the vanes. A portion of the axial length of the radially inner sealing ring comprises a radially inner wall and a radially outer wall spaced radially from the radially inner wall to define a chamber. The radially inner wall has apertures to supply coolant into the chamber and the radially outer wall has effusion cooling apertures to supply coolant onto the surface of the radially outer wall.

Abstract: A fuel spray nozzle is provided for atomising liquid in a gas, the nozzle comprising: a gas passage; a liquid passage; a prefilming surface positioned downstream of the liquid passage and the gas passage, and configured to receive liquid from the liquid passage and to receive gas from the gas passage; wherein the liquid passage terminates at an exit orifice upstream of the prefilming surface; and wherein the gas passage contains a flow perturbator upstream of the liquid passage exit orifice, to increase the turbulence of gas passing from the gas passage to the prefilming surface.

Abstract: A tip clearance control (TCC) system 100 is provided to control the gap 176 between the tips of turbine blades 172 of a gas turbine engine 10 and the casing 174 within which they rotate. The inlet 120 to the TCC system is provided in a bifurcation panel 110 that extends across a bypass duct 22 of the gas turbine engine 10. The inlet is provided on a first major surface 112 of the bifurcation panel 110 that is defined such that the direction (R) that points through the bifurcation panel from the first major surface 112 to a second major surface 114 corresponds to the fan rotation direction. This provides a particularly effective and/or efficient TCC system 100.

Abstract: A system for minimising the formation of ice crystals in the fuel system pipework of a gas turbine engine. The system includes an ultrasonic transmitter located within a wall of a pipe in the fuel system through which fuel passes. The system also includes a controller that causes said ultrasonic transmitter to produce ultrasonic waves when the temperature of the fuel is sufficiently low for any water in the fuel to freeze, the ultrasonic waves being sufficient to break particles of ice contained in the fuel that passes through the pipe. A method of minimising the formation of ice crystals in the fuel system pipework of a gas turbine engine is also disclosed.

Abstract: A switched reluctance machine is provided, the machine includes a stator having a circumferential row of magnetically energizable teeth carrying coil windings. The machine further includes a coaxial variable reluctance rotor having a circumferential row of magnetisable poles, the reluctance of the rotor-to-stator magnetic flux path varying with rotor position. Each pole has first and second side flanks and a tip surface, one of the side flanks forming a leading edge surface of the pole, the other side flank forming an opposite, trailing edge surface of the pole, and the tip surface arcing from the leading edge surface to the trailing edge surface and facing the stator to form an air gap to the teeth of the stator. Also each pole contains one or more cavities.

Abstract: A method of assembling an annular combustion chamber assembly comprises positioning a plurality of first combustion chamber segments circumferentially side by side to form an annulus and removably securing the downstream end of each first combustion chamber segment to a first ring to form a first assembly. The method comprises positioning a plurality of second combustion chamber segments circumferentially side by side to form an annulus, removably securing the downstream end of each second combustion chamber segment to a second ring to form a second assembly. The method comprises removably securing the upstream end of each second combustion chamber segment of the second assembly to an annular upstream end wall. The method comprises inserting the second assembly into the first assembly and then removably securing the upstream end of each first combustion chamber segment of the first assembly to the annular upstream end wall to form the annular combustion chamber assembly.

Abstract: Apparatus for controlling velocity of aircraft during landing roll-out and/or taxiing, the apparatus comprising: a generator for absorbing kinetic energy from a landing gear of the aircraft to generate electrical energy; and a component of the aircraft for receiving and consuming electrical energy from the generator, the generator and the component being electrically connected to one another without an intervening electrical energy storage device.

Abstract: A nickel based alloy composition including, by atomic per cent, between 12 and 15% of elements from the group of aluminium, titanium, tantalum and niobium, between 0.8 and 1.8% tungsten, between 5 and 8% aluminium, at least 28% of elements from the group of molybdenum, tungsten, chromium and cobalt, wherein the atomic ratio of the sum of titanium, tantalum and niobium to aluminium is between 0.6 and 1.1, and wherein the composition includes less than 1.5% molybdenum, the balance being nickel save for incidental impurities.

Abstract: The present invention provides a method of making a metallic shield for shielding an edge of a component e.g. an aerofoil component such as a gas turbine blade or vane. The method comprises: providing a mandrel with a mandrel edge having a shape conforming to the shape of the component edge, wherein the mandrel edge comprises a surface having a width, W1; providing a metal wire having a substantially triangular radial cross-section with a base surface of width W2, wherein W2 equals W1; aligning and affixing the base surface of the wire on the surface of the mandrel edge; electro-depositing metal onto the mandrel and metal wire to form the metallic shield comprising the deposited metal and the metal wire; and removing the metallic shield from the mandrel.

Abstract: A method of reinforcing a fiber reinforced resin matrix composite structure, the method including: providing a composite structure having one or more holes formed therein, and feeding a reinforcement rod through a passageway having one or more walls defining an inlet to the passageway. The inlet to the passageway is dimensioned to prevent a rod having a maximum cross sectional dimension greater than a predetermined dimension from passing through the passageway. In the event of the rod impacting the wall defining the inlet to the passageway, operation of the rig is stopped. In the event of the rod passing through the passageway without impacting the wall defining the inlet to the passageway, a portion of the rod is inserted into a hole in a composite structure.

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.

Abstract: A gear comprising: a hub defining an outer surface, the hub comprising a ceramic material; and a first ring defining an inner surface and an outer surface, the inner surface of the first ring being coupled to the outer surface of the hub via a bond, the outer surface of the first ring defining a plurality of gear teeth, and the first ring comprising a metallic material.