Abstract: A composite flange element formed on a casing comprising a portion of a component and an adjoining flange portion, the composite having a ply layup comprising: one or more first plies extending over both the portion of the component and the flange portion; one or more second plies extending over the flange portion; and one or more third plies extending over the portion of the component; wherein the first, second and third plies are selected to provide desirable properties for the portions over which they extend. The flange element so formed consists of one or more load spreading features associated with the flange portion and one or more stress reduction features in the region where the portion of the component meets the flange portion. Preferably the plies are arranged or interleaved such that they run up the case and flange without distortion.

Abstract: A containment system 20 is described, e.g. for a gas turbine engine. The system 20 includes a containment casing 22 and may also include a face sheet 30 and a backing layer 32. The casing 22 includes a matrix material 24 and regions 26 formed of a second material different to the matrix material. Fibres 28 are also included, formed of a third material.

Abstract: A method is provided that enables the identification of a region of a layered composite component that a test signal has passed through. The method comprises applying a label to the test signal as it passes through the region, such that the label applied to the test signal is indicative of the said region. The labelled signal can be identified by a sensor, even if the signal is of low strength and amongst a lot of noise, such as that caused by the multiple reflections and refractions that take place when an acoustic signal passes through a layered composite component.

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

Abstract: A method of manufacturing a component comprising a hollow axisymmetric body formed of a cured composite material is provided. The method includes the step of positioning the constituents of the composite material in a mould corresponding to at least a portion of the hollow axisymmetric body. The composite material is cured into a composite part, and residual stresses are generated in the composite material during the curing by controlling the temperature of the constituents in the mould. When the composite part is released from the mould, relaxation of the residual stresses causes one or more regions of the composite part to deflect in the radial direction of the hollow axisymmetric body, the position of greatest radial deflection of each region having a radial deflection which is at least 0.1% of the greatest diameter of the hollow axisymmetric body.

Abstract: A mould for a composite component having a fillet joint, the mould including a mould body having a cavity within which the component is formed, and an insert mating with the mould body and including a forming surface against which, in use, the fillet of the component fillet joint is formed. The insert may be thermally insulating and highly rigid. A method of forming a composite component having a fillet joint using a mould, the method including forcing a surface of the fillet region of the component to adopt a non constant radius of curvature while the component is inside the mould during the cure process is also provided. Additionally, a method of forming a composite component having a fillet joint that is defined between first and second legs the method including placing the legs of the component in bending during cure of the component is described.

Abstract: A composite material and method for manufacturing the material, the material including: a plurality of plies layered one on top of the other; and one or more through-thickness fibres which join one or more of the plurality of plies to one another; wherein the one or more through-thickness fibres form a boundary which delineates one or more discrete regions of the material.

Abstract: A fibre cutting device and method for cutting one or more exposed portions of a through-thickness fibre in a surface of a composite material, the device including: a cutting element for cutting the one or more exposed portions; a suction device which is operable to extract the cut exposed portions from the surface of the composite material; and a storage device for receiving the cut exposed portions.

Abstract: Aerofoils (22) of a gas turbine engine are provided with a coating (34) or filler (44) of viscoelastic material. As ice accretes on the aerofoils (22) during operation, the resulting aerodynamic stability imbalance induces vibration in the aerofoils (22). The viscoelastic material (34, 44) damps this vibration, and in so doing generates heat, which melts the ice away from the aerofoils (22). Heat-conducting members conduct the heat to regions of the component in which ice accretion is to be prevented. Alternative embodiments are described in which the pseudoelastic behaviour of a shape memory alloy (56), or eddy currents arising from the rotor blades' rotation in an axisymmetric magnetic field, are used as sources of heat.

Abstract: A composite blade (26) comprises a three-dimensional arrangement of reinforcing fibres (58) and a matrix material (60) infiltrated around the three-dimensional arrangement of woven reinforcing fibres (60). The three-dimensional arrangement of woven reinforcing fibres (58) defines a plurality of cavities (56) within the aerofoil (28). The composite blade (26) comprises an aerofoil portion (38) and a root portion (36). The aerofoil portion (38) comprises a leading edge (44), a trailing edge (46), a concave pressure surface wall (50), a convex suction surface wall (52) and a tip (48). The aerofoil portion (36) comprises a plurality of webs (54) extending between, and being secured to, the concave pressure surface wall (50) and the convex suction surface wall (52) to produce a Warren girder structure. The three-dimensional arrangement of woven reinforcing fibres (58) are arranged to produce the concave pressure surface wall (50), the convex suction surface wall (52) and the plurality of webs (54).

Abstract: A composite blade (26) comprises a three-dimensional arrangement of reinforcing fibres (58) and a matrix material (60) infiltrated around the three-dimensional arrangement of woven reinforcing fibres (60). The three-dimensional arrangement of woven reinforcing fibres (58) defines a plurality of cavities (56) within the aerofoil (28). The composite blade (26) comprises an aerofoil portion (38) and a root portion (36). The aerofoil portion (38) comprises a leading edge (44), a trailing edge (46), a concave pressure surface wall (50), a convex suction surface wall (52) and a tip (48). The aerofoil portion (36) comprises a plurality of webs (54) extending between, and being secured to, the concave pressure surface wall (50) and the convex suction surface wall (52) to produce a Warren girder structure. The three-dimensional arrangement of woven reinforcing fibres (58) are arranged to produce the concave pressure surface wall (50), the convex suction surface wall (52) and the plurality of webs (54).

Abstract: Aerofoils (22) of a gas turbine engine are provided with a coating (34) or filler (44) of viscoelastic material. As ice accretes on the aerofoils (22) during operation, the resulting aerodynamic stability imbalance induces vibration in the aerofoils (22). The viscoelastic material (34, 44) damps this vibration, and in so doing generates heat, which melts the ice away from the aerofoils (22). Heat-conducting members conduct the heat to regions of the component in which ice accretion is to be prevented. Alternative embodiments are described in which the pseudoelastic behaviour of a shape memory alloy (56), or eddy currents arising from the rotor blades' rotation in an axisymmetric magnetic field, are used as sources of heat.

Abstract: A strain gauge assembly 2 comprising: a plurality of strain gauges 4, 6 arranged in a side-by-side relationship, each having a first end 10 and a second end 12; wherein the first and/or second end 10, 12 of one of the plurality of strain gauges 4, 6 is offset from a corresponding first and/or second end 10, 12 of another of the plurality of strain gauges 4, 6.

Abstract: A layered composite component 100, 200 is disclosed comprising a plurality of plies of reinforcement fibres embedded in a matrix material. The component comprises a fuse region 134, 144, 157, 233, 244 in which an initiating feature 180, 190, 155 is operable to initiate delamination of plies within the fuse region 134, 144, 157, 233, 244 such that the fuse region 134, 144, 157, 233, 244 delaminates above a predetermined load condition. The component 100, 200 is operable to transition from a rigid behaviour regime to a resilient behaviour regime on occurrence of delamination within the fuse region 134, 144, 157, 233, 244.

Abstract: A composite material comprises plies 2, 4, 6, 8, which may be prepregs. Film elements 10 are provided between at least two adjacent ones of the plies 2, 4, 6, 8. The film elements are made from a release material which inhibits adhesion between regions of the resin on opposite sides of the respective film elements, and are randomly distributed. The film elements 10 provide delamination sites within the component. The component may be a test piece for assessing the delamination behaviour of the material. Alternatively, the delamination behaviour caused by the film elements 10 may have a useful purpose, for example to control regions of a component in which delamination occurs following events such as a bird strike on a fan blade of a gas turbine engine.

Abstract: A process for manufacturing a layered composite component includes introducing a plurality of micro cracks into a predetermined region of the component in order to provide an increase in strength of the component at certain critical regions of high loading within the component. The micro cracks may be introduced using mechanical means, a pre conditioning load, or by manipulation of process parameters, material properties or stacking sequences.

Abstract: A containment casing is provided comprising a radially outer layer; an intermediate energy absorbing layer, and a retention framework for retaining an aerofoil blade in the event of detachment of whole or part of the blade. The retention framework is distributed axially and circumferentially along the casing and extends substantially radially inwardly at least partially through the intermediate energy absorbing layer.

Abstract: A vane assembly for a gas turbine engine, the assembly including a plurality of vanes 14 with an asymmetric organic metal/ceramic matrix composite lay up, such that when the vanes 14 are placed in tension or compression the profile of the vane 14 will alter by twisting or untwisting. Other materials may be provided for the vanes 14, and/or other methods may be used for placing the vanes 14 in tension or compression to provide positive changes to the profile thereof.

Abstract: A lightweight component in the form of an outlet guide vane for a gas turbine engine and a method of manufacturing such lightweight components are provided. The vane is filled with hollow metal ellipsoid members in a solid metal outer member. The arrangement of hollow metal ellipsoid members allows directionally tailored stiffness or strength.

Abstract: A component, for example an aerofoil vane or other gas path structure in a gas turbine machine such as a gas turbine engine, in which an insulated electrical conductor path (2) is embedded structurally integrally, and in which electromagnetic shielding (1, 2, 20, 22), for the embedded electrical conductor path is provided structurally integrally.