Abstract: An annulus filler is provided for mounting to a rotor disc of a gas turbine engine. The filler comprises a lid for bridging the gap between two adjacent blades attached to the rotor disc. The filler further comprises forward and rearward hook members projecting below the lid and configured to respectively engage with forward and rearward outwardly directed hooks provided on the rotor disc when the annulus filler is translated in a generally rearwards direction on mounting of the annulus filler to the rotor disc. The filler can further comprise a lip extending from the rear of the lid, the lip being configured to fit beneath a gas-washed element located rearwards of the annulus filler. One or more guide ramps then extend from the rear of the lip, the or each guide ramp being angled down relative to the lip to encourage the lip to slide beneath said fairing when the annulus filler is translated in the generally rearwards direction.

Abstract: A component, such as a blade, vane or combustor wall of a gas turbine engine comprises two walls defining a coolant passage and has an array of pedestals extending between the two walls for heat removal. Each pedestal changes in cross-section along its length. Alternate rows of pedestals are arranged such that their larger cross-sectional area is adjacent one wall then the other. When a coolant flows through the passage it is forced to flow between one wall and the other wall so as to increase turbulence and hence mixing for a more even coolant temperature. The array of pedestals can also be used to tailor the individual heat loads on each wall independently and has the ability to use differing levels of blockage to counter adverse pressure gradients along successive rows of pedestals.

Abstract: A pressure actuated tip clearance system for a shroud structure of a gas turbine engine rotary stage includes an annular plenum chamber formed between an annular arrangement of a plurality of shroud liners on the inner circumference of the chamber and a generally cylindrical casing on the radially outer side, and, in use, a hot gas stream radially inwards of the shroud liners. Each shroud liner comprises a hollow box section comprising upstream and downstream walls, radially inner and outer walls, and side walls, the downstream and radially inner and outer walls being closed, the upstream wall having an air inlet aperture, and at least one of the side walls having at least one outlet aperture. The inlet aperture is in flow communication with a source of high pressure air at a pressure higher than that of the gas stream.

Abstract: A male or female element (12, 2) for a conic coupling, the element comprising: an end surface (16, 6) and two angled side surfaces (18, 8) disposed either side of the end surface (16, 6); wherein at least one of the side surfaces (18, 8) has a non-planar profile.

Abstract: A male or female element (12, 2) for a conic coupling, the element comprising: an end surface (16, 6) and two angled side surfaces (18, 8) disposed either side of the end surface (16, 6); wherein a coating is applied to at least one of the side surfaces, the coating having a varying coefficient of friction across the surface.

Abstract: A method of producing an oxide dispersion strengthened nickel-base superalloy, comprising introducing an oxide dispersoid material into a plasma gun of a plasma spray apparatus, where it is sublimed and turned to vapour; and introducing a nickel-base superalloy material into the plasma spray apparatus at a cooler location, downstream of the plasma gun, such that the oxide dispersoid material condenses on the superalloy material to produce the oxide dispersion strengthened nickel-base superalloy.

Abstract: A boundary layer energiser (20) for energising a boundary layer flow (33) over a surface (22), the boundary layer energiser (20) comprising one or more passages (24) terminating in one or more respective holes (26) provided on the surface (22), wherein the one or more passages (24) comprises at least one fin (25) configured so as to generate a vortex flow in the one or more passages (24) such that, when in use, a fluid emanating from the one or more passages (24) flows in a vortex.

Abstract: A gas turbine engine comprising a nacelle and a breather duct. The breather duct provides communication between a component of the engine within the nacelle and the exterior of the nacelle and opens at an exhaust port on the external surface of the nacelle. There is at least one clean air outlet bounding the exhaust port and arranged to expel, in use, a vortical flow of clean air to form an aerodynamic barrier between the external surface of the nacelle and a flow of breather air expelled from the exhaust port.

Abstract: A boundary layer energiser (20) for energising a boundary layer flow over a surface (22), the boundary layer energiser (20) comprising a plurality of passages (24), each passage terminating in a respective hole (26) provided on the surface (22), the holes being arranged in a cluster (23) on the surface, wherein the plurality of passages are angled with respect to one another at the surface such that, when in use, a vortex (24) is formed by a fluid flowing through the plurality of passages.

Abstract: Fluid release apparatus (26) for the release of fluid from the interior of a rotating and/or pressurized body (16;18;102;104), the apparatus comprising a core member (28) having a chamber wall (44) disposed thereabout so as to define an open chamber (30) about said core member. The apparatus further comprises a chamber closure (48) mounted about said core member (28), said closure being resiliently deformable between an at rest state in which the closure contacts the chamber wall (44) and an actuated state in which the closure is spaced there-from, wherein the chamber has an aperture (32;35;55) for receiving liquid phase fluid from the interior of said rotating body. Upon build up of sufficient liquid in said chamber (30) the closure (48) moves to said actuated state so as to release liquid from the interior of the body.

Abstract: Removal of a soluble part of a core allows provision of relatively complex three-dimensional geometric shapes in a component. Unfortunately previous leaching arrangements have been relatively erratic such that it is difficult to provide consistency of component form over a batch of components formed. By providing a tank combination which comprises typically a number of dip tanks, it is possible to provide consistency in terms of leach solution efficiency upon cores. Generally leaching solution held within the tanks is adjusted through either a prior calibrating adjustment tank or through agitation or other means in the tank such that the effects of saturation ageing of the leaching solution can be homogenised across all cores subject to the leaching process.

Abstract: A displacement amplifier 10 comprises first and second longitudinally extending elements 12, 22, each having a free portion 14, 24 and a constrained portion 16, 26. The constrained portions 16, 26 of the first and second elements 12, 22 are constrained to move together such that longitudinal displacement of a free portion 14 of one element with respect to the free portion 24 of the other element results in an amplified displacement of the constrained portions 16, 26 relative to a resting position.

Abstract: A surface treatment device (10) is provided for applying a surface treatment medium to an article (60). The device comprises an application member (12, 14) for holding and applying the surface treatment medium to the article. The device has a securing arrangement (40) for securing the application member to the article.

Abstract: A method, for detecting overpressure inside a compartment associated with a gas turbine nacelle, includes: a) measuring one or more predetermined operating conditions nominally affecting the temperature inside the compartment; b) determining a reference temperature for the compartment, corresponding to a nominal reference zone pressure, on the basis of the measured operating condition or conditions; c) measuring the actual temperature inside the compartment, corresponding to the actual compartment pressure; and d) using the measured actual temperature in the compartment and the determined reference temperature for the compartment to assess the gas pressure inside the compartment. A system for assessing the gas pressure inside the compartment caused by an associated pressure system failure includes using a quantified mass leakage flow to assess gas pressure inside the compartment.

Abstract: An apparatus for measuring erosion of an edge of an aerofoil portion of a fan blade includes a laser source to direct light at a position on the edge of the fan blade. A detector detects light reflected from the position on the edge of the fan blade. A computer measures the distance from the laser source to points on the edge of the fan blade to produce a profile of the edge in terms of x and z coordinates. The computer calculates the distances of points on the edge of the fan blade from a centroid of the profile. The computer detects peaks in the distance of points from the centroid of the profile. The computer selects two peaks with the greatest distance therebetween. The computer determines the distance between the two peaks and compares the determined distance and a predetermined distance to determine if there is unacceptable erosion.

Abstract: A method of forging a nickel base superalloy comprising providing a nickel base superalloy preform (40) with a first predetermined shape, the nickel base superalloy preform having been produced by powder metallurgy. The nickel base superalloy preform (40) is forged to produce a nickel base superalloy forged component (50) with a second predetermined shape. The first predetermined shape and the second predetermined shape are arranged such that the effective strain at the end of the forging is less than 1. The nickel base superalloy forged component (50) is given a supersolvus heat treatment to produce a large grain size in the nickel base superalloy forged component (50).

Abstract: A combustor is provided for mounting a fuel injector to a gas turbine engine. The combustor comprises an engine casing having an aperture formed therein. The system further comprises a fuel injector having a flange for mounting the fuel injector to the casing at the aperture so that the fuel injector extends into the engine. The flange is dismountably sealed to an inner side of the casing. The aperture and flange are configured so that, when dismounted, the fuel injector can be rotated into an orientation relative to the aperture which allows the flange to pass though the aperture and the fuel injector to be withdrawn from the casing.

Abstract: A system is provided for mounting fuel injectors to a gas turbine engine. The system comprises an engine casing having a plurality of apertures formed therein. The system further comprises a plurality of first and second fuel injectors. Each fuel injector has a respective flange for mounting the fuel injector to the casing at a respective aperture so that the fuel injector extends into the engine. The first fuel injectors have flanges which are dismountably sealed to an inner side of the casing. The second fuel injectors have flanges which are dismountably sealed to an outer side of the casing. The flanges of the first fuel injectors are configured to allow them to pass through the apertures of the second fuel injectors. A first fuel injector can be dismounted from the casing and withdrawn therefrom through the aperture of a dismounted second fuel injector.

Abstract: A system for mounting a fuel injector to an engine casing with an aperture includes a fuel injector having a flange for mounting to the casing at the aperture. The flange allows it to pass through the aperture. An intermediate ring mediates the mounting of the flange to the casing at the aperture. The intermediate ring is inside the casing at the aperture and defines an opening from which the fuel injector extends into the engine with the flange positioned inward of the ring. The flange is dismountably sealed to an inner side of the ring, and the ring is dismountably sealed to an inner side of the casing to mount the fuel injector to the casing. On dismounting the flange from the ring, the fuel injector can be displaced to allow the ring, to be moved away from the aperture. The fuel injector can then be withdrawn.

Abstract: A ducted fan gas turbine engine aerofoil is made by electron beam welding together at least two metal sheets (10) and (12) and electron beam welding that sub assembly via an end to a root that has been manufactured in a separate operation, and then heating the whole to a temperature that will convert the electron beam welds to diffusion bonds.