Abstract: A panel for lining a gas turbine engine fan casing includes a honeycomb core sandwiched between a backing skin and an outer skin. The backing skin is attached to an inner surface of the casing such that the outer skin forms a radially inward facing surface of a fan duct of the engine. The panel is joined along its sides to similar neighbouring panels such that, in use, the joined panels form a row of panels along the inner surface of the casing. The outer skin or the backing skin includes two face sheets bonded on top of each other, which are arranged such that their edges along each panel side that joins to a neighbouring panel are staggered in the direction of the row. The interfaces between the abutting face sheet edges are therefore correspondingly staggered in the direction of the row.

Abstract: The present invention provides a gas turbine engine comprising a tubular containment casing surrounding a rotary fan blade assembly. The radially outer perimeter (and optionally the radially inner perimeter) of the radial cross-sectional profile of the containment casing is non-circular e.g. polygonal, corrugated, fluted or wavy.

Abstract: A fan casing assembly comprising an annular casing member for circumferentially surrounding a fan, and an annular fan track liner comprising a plurality of circumferentially adjacent liner panels. The annular fan track liner is positioned radially inward of the annular casing member. A tilting arrangement is provided and configured to selectively circumferentially tilt one or more of the liner panels, such that, in the event of a fan blade being released from a fan, the tilting arrangement can initiate tilting of one or more liner panels so that the liner panels can trim one or more of the remaining fan blades.

Abstract: A panel for lining a gas turbine engine fan casing includes a honeycomb core sandwiched between a backing skin and an outer skin. The backing skin is attached to an inner surface of the casing such that the outer skin forms a radially inward facing surface of a fan duct of the engine. The panel is joined along its sides to similar neighbouring panels such that, in use, the joined panels form a row of panels along the inner surface of the casing. The outer skin or the backing skin includes two face sheets bonded on top of each other, which are arranged such that their edges along each panel side that joins to a neighbouring panel are staggered in the direction of the row. The interfaces between the abutting face sheet edges are therefore correspondingly staggered in the direction of the row.

Abstract: A fan casing assembly comprising an annular casing member for circumferentially surrounding a fan, and an annular fan track liner comprising a plurality of circumferentially adjacent liner panels. The annular fan track liner is positioned radially inward of the annular casing member. A tilting arrangement is provided and configured to selectively circumferentially tilt one or more of the liner panels, such that, in the event of a fan blade being released from a fan, the tilting arrangement can initiate tilting of one or more liner panels so that the liner panels can trim one or more of the remaining fan blades.

Abstract: The present invention provides a gas turbine engine comprising a tubular containment casing surrounding a rotary fan blade assembly. The radially outer perimeter (and optionally the radially inner perimeter) of the radial cross-sectional profile of the containment casing is non-circular e.g. polygonal, corrugated, fluted or wavy.

Abstract: A method of forming a component includes defining an identification pattern, defining one or more scanning parameters and/or one or more heating parameters, depositing a sinterable material on a substrate and scanning a heat source of the deposited sinterable material to thereby selectively sinter the material to the substrate to produce a sintered layer having the identification pattern. The sintered layer includes first and second regions, and the method includes sintering the first region using one or more scanning parameters and/or one or more heating parameters having a first value, and sintering the second region using one or more scanning parameters and/or one or more heating parameters having a second value to thereby produce the identification pattern including a contrast between the first and second regions.

Abstract: This invention relates to a stator aerofoil for a gas turbine engine, comprising: a body portion; at least one panel which forms at least part of one of either the pressure or section surface of the aerofoil; at least one internal chamber which is bounded by the body portion and panel; and, at least one elastomeric component within the at least one internal chamber, wherein the elastomeric component at least partially defines a cell within the internal chamber, wherein the cell is in fluid communication with the exterior of the aerofoil.

Abstract: A mandrel includes a support structure; main sectors carried by the support structure and each having a main forming surface; and intermediate sectors carried by the support structure and each having an intermediate forming surface. The main sectors and intermediate sectors are alternately arranged around the axis. The main sectors are moveable relative to the support structure between a radially inner main retracted position, and a radially outer main forming position. The intermediate sectors are moveable relative to the support structure between a radially inner intermediate retracted position, and a radially outer intermediate forming position. The mandrel further includes drive means for moving the main sectors and intermediate sectors from their respective retracted positions to their respective forming positions. The drive means is configured such that each main sector reaches its main forming position before the adjacent intermediate sectors reach their respective intermediate forming positions.

Abstract: Apparatus 200 for forming a composite component is disclosed, the apparatus comprising: a deformable mold tool 204 that at least partially defines a mold cavity 212; an interface layer 222 formed on an external surface of the mold tool 204; and at least two peristaltic actuators 206, mounted to contact the interface layer 222 and to generate a peristaltic pressure wave in the mold tool 204. The peristaltic actuators comprise rollers 206, pistons 306 or bi-metallic blocks 406. A method of forming a composite component is also disclosed, the method employing a mold tool 204 that at least partially defines a mold cavity 212. The method comprises: placing reinforcing fibers into the mold cavity 212, forcing a matrix material into the mold cavity 212 through the reinforcing fibers, and generating a peristaltic pressure wave within the mold tool 204 during forcing of the matrix material into the mold cavity 212.

Abstract: An annulus filler system bridges the gap between two adjacent blades attached to a rim of the rotor disc of a gas turbine engine. The system includes an annulus filler having a lid which extends between the adjacent blades and defines an airflow surface for air being drawn through the engine. The filler also has a support body extending beneath the lid and terminating in an elongate foot which, in use, extends along a groove provided in the rim of the disc. The groove has a neck which prevents withdrawal of the foot through the neck in a radially outward direction of the disc. The system further includes a sleeve which, after installation of the filler, is slidably locatable into a gap between the foot and sides of the groove.

Abstract: A method of imparting deep compressive residual stress to an aerofoil, the method involves determining a stress map for the aerofoil for stress generated by Foreign Object (FOD) impact, a stress map for the aerofoil of high cycle fatigue, a stress map for the aerofoil of low cycle fatigue and determining a combined stress map by combining the stress map of FOD impact, the stress map of high cycle fatigue and the stress map of low cycle fatigue. A zone is then defined on the aerofoil from the combined stress map for receiving the deep residual compressive stress, and compressive residual stress imparted to the defined zone.

Abstract: A fluid management apparatus and method, the apparatus including: a fluid conduit for the passage of a dispersion containing particulate matter; a laser arranged to provide laser light inside the fluid conduit; wherein, in use, the laser light heats the particulate matter sufficiently to generate incandescence; one or more sensors for detecting the incandescence of the particulate matter so as to determine a characteristic of the dispersion; and an electromagnetic wave generator operable to provide electromagnetic waves inside the fluid conduit at a position downstream of the laser light so as to vaporize at least a portion of the dispersion.

Abstract: An assembly for the removal of contaminants from a contact zone between respective contact surfaces 6, 8 of a first component 2 and a second component 4. At least one of the contact surfaces has a low friction element 10 provided with grooves 12 which extend through the contact zone such that, in operation, a pressure difference across the contact zone causes contaminants entering the grooves 12 to be expelled along the grooves 12 from the contact zone.

Abstract: A turbomachine casing assembly including a first casing element locatable radially outward of one or more rotating aerofoil elements of a turbomachine, a second casing element located radially distal to the first casing element is provided. A void is provided between a radially proximal face of the first casing element and a radially proximal face of the second casing element, at the first end of the first casing element. The first end of the first casing element is located against a corresponding first end of the second casing element by an energy absorbing element, with the first casing element; including a cantilever. The cantilever is arranged in a region extending along the first casing element from the first end; and having a pivot axis at a mid portion thereof.

Abstract: An annulus filler assembly for a rotor of a turbomachine, the assembly comprising: an annulus lid (56) having a radially outwardly facing surface (58) for forming an inner wall of a flow annulus of the rotor and a radially inwardly facing surface; and a frame for supporting the annulus lid (40), the frame being mountable to a disc of the rotor such that the annulus lid (56) is spaced away from the disc, wherein the frame (40) comprises a connection portion (46) which, in use, passes through an aperture (60) in the annulus lid (56) from the radially inwardly facing surface towards the radially outwardly facing surface (58) such that at least a portion of the connection portion (46) is visible from the radially outwardly facing surface (58); the assembly further comprising a locking element (68) which locks the connection portion (46) to the annulus lid (56) via the visible portion of the connection portion (46).

Abstract: A joint assembly includes a first component with a flange containing a plurality of spaced through-holes; a second component with a joining portion having an end surface which faces the flange and containing a plurality of spaced elongate cavities; and a plurality of fixation devices. Each cavity extends into the joining portion from the end surface in alignment with one of the through-holes. Each fixation device includes: a bolt which extends through a respective through-hole and the corresponding cavity, and includes a nut which is embedded within the joining portion at the corresponding cavity to receive the bolt such that on tightening the nut and bolt a clamping force is exerted across the flange and the end surface to join the first and second components together. Under axial loads less than that causing bolt failure, each fixation device is deformable to remove the clamping force.

Abstract: A component for use in a gas turbine engine comprises an indicator operable to provide a visual indication when strain over a threshold value has been experienced by the component by transitioning from a rest state to an indication state on occurrence of strain above the threshold value. Also disclosed is a method of detecting foreign object damage in a component of a gas turbine engine, comprising: identifying at least one of a component or a region of a component that is subject to increased strain during a foreign object impact; and mounting an indicator on the identified component or region of a component, the indicator providing a visual indication when increased strain has been experienced.

Abstract: The present invention relates to a bearing assembly (1) for a rotatable shaft (2). The bearing assembly comprises a bearing housing (10); a bearing (20) located within the bearing housing (10) and arranged in use to receive a rotatable shaft (2); and an elongate member (14) that couples the bearing (20) to the bearing housing (10) for damping vibrations of the rotatable shaft (2). The elongate member (14) comprises a piezoelectric element (40) which may be actuated so as to exert a force on the elongate member. This arrangement allows the bearing assembly to damp vibrations of the rotatable shaft.

Abstract: The present invention relates to a bearing assembly (1) for a rotatable shaft (2). The bearing assembly comprises a bearing housing (10); a bearing (20) located within the bearing housing (10) and arranged in use to receive a rotatable shaft (2); and a damper (14) that couples the bearing (20) to the bearing housing (10) for damping vibrations of the rotatable shaft (2). The damper (14) comprises a shape memory alloy (40). This arrangement allows the bearing assembly to damp vibrations of the rotatable shaft.