Abstract: A composite blade (28) comprises a root portion (38) and an aerofoil portion (40). The aerofoil portion (40) has a tip (50) remote from the root portion (38). The aerofoil portion (40) has a leading edge (42), a trailing edge (44), a suction surface (46) extending from the leading edge (42) to the trailing edge (44) and a pressure surface (48) extending from the leading edge (42) to the trailing edge (44). The composite blade (28) comprises reinforcing fibers (62) arranged in a matrix material (63). The tip (50) of the aerofoil portion (40) is arranged to define a recess (66), an insert (68) is arranged in the recess (66) and the insert (68) comprises an abradable material.

Abstract: A seal and method of making seals utilises folds in order to form creases in a length of material such that the creases reinforce individual seal elements or surfaces whilst perforations or slots in a seal edge allow air leakage from one side of the seal element or surface to the other. Thus, air pressurisation can take place between spaced seal elements or surfaces and this allows air flotation or riding of a rotating component relative to a static housing component. The perforations are typically graduated from the seal edge in order to provide the best pressurisation and air-riding effect for proximity or gap control between the sealing edge and the rotating component surface. The slots effectively provide flexibility to the sealing edge such that the seal component emulates a brush seal.

Abstract: A computer system programmed to process a large data set includes means for analysing the data set. The means for analysing the data set typically comprises a means for finite element analysis. The computer system also includes means for applying a data compression technique to the analysed data set such that the compressed analysed data set has high fidelity in regions of interest and has lower fidelity in regions of lesser interest. The data compression technique comprises the use of a wavelet compression technique. The computer system comprises a high specification server and a local workstation. This compressed analysed data set is easily managed by the local workstation to produce different cross-sectional displays without reusing the high capability server. The analysed data set for example comprises a 4D data set of a fan blade containment analysis of a casing.

Abstract: A guide vane arrangement 20 for a gas turbine engine (10, FIG. 1) includes a vane member 21 extending between inner and outer platforms 22, 24 which are respectively mounted on inner and outer mounting members 42, 34. One of the inner and outer platforms 22, 24 includes a resilient that abuts the respective inner or outer platform 22, 24 to permit relative movement between the inner or outer platform 22, 24 and the respective inner or outer mounting member 42, 34.

Abstract: A seal and method of making seals utilizes folds in order to form creases in a length of material such that the creases reinforce individual seal elements or surfaces while perforations or slots in a seal edge allow air leakage from one side of the seal element or surface to the other. Thus, air pressurization can take place between spaced seal elements or surfaces and this allows air flotation or riding of a rotating component relative to a static housing component. The perforations are typically graduated from the seal edge in order to provide the best pressurization and air-riding effect for proximity or gap control between the sealing edge and the rotating component surface. The slots effectively provide flexibility to the sealing edge such that the seal component emulates a brush seal.

Abstract: Composite aerofoils for gas turbine engines are commonly provided with a metal protection strip along the leading edge, to prevent erosion of the leading edge in use and to protect against impacts from foreign bodies. A problem with such strips is that they can cause serious damage to other parts of the engine if they become detached from the aerofoil. The invention provides an aerofoil having such a protection strip, characterised in that the protection strip includes one or more weakening features to reduce the ability of the protection member to withstand a compressive force applied along its length. The weakening features encourage the protection member to break up under impact, or if it becomes detached from the aerofoil, so that damage to other parts of the engine is minimised.

Abstract: Previous testing of such test pieces as fan blades utilized in gas turbine engines has involved application of paint markings to identify fragments upon fragmentation of the test piece. Such fragmentation causes debris which will obscure paint markings and the paint itself may create clouds of dust obscuring visual images of the test piece under test conditions. By applying an image pattern comprising a number of lingering image patches, typically in the form of thermal patches upon the test components, and arranging the test components to be within an enclosure which is evacuated it will be appreciated that these image patterns linger and can be viewed by an appropriate monitor over a period at least of the test conditions. Utilization of extra visual spectrum radiation images extends the period when clear images of the test piece under test conditions can be seen and analyzed.

Abstract: A stage of fan aerofoils (10) lies within a fan cowl (12). The fan duct (16) is defined in part by a hard casing (14) that in turn surrounds aerofoils (10). Hard casing (14) includes wedge members (26) that fill the annular gap between ring (14) and an outer ring (20). In the event of an aerofoil (10) breaking off, the hard ring (14) and wedges (26) absorb sufficient of the kinetic energy expended by the broken aerofoil (10), as to prevent it passing through outer ring (20) on to the fan cowl (12).

Abstract: The ducted fan (12) of a ducted fan gas turbine engine (10) is surrounded by a compartment (26), which is packed with bristles (38). Should a blade of the fan stage (12) disintegrate during operation of engine (10), the resulting fragments, if they pass through an abradeable honeycomb structure (32), will collide with the bristles (38), which will absorb the energy exerted by the fragments, and thus prevent puncturing of the outer wall (18) of the fan duct structure.

Abstract: Blades for gas turbine engines which are formed from composite materials have problems with respect of resistance to impacts such as bird strikes. Previous blades formed from metals had some ductility towards the trailing edge which could accommodate the whiplash effects of impacts. With regard to composite materials such ductility is not present. By providing projections 32 which act as propagation wave trips as well as high intensity reflectors 36 it is possible to limit the whiplash at the edge 31 resulting in damage. Typically a cladding cap 38 is provided which also may be formed from a metal to allow some greater uniformity with respect to mass per length despite the tapering of the blade. Furthermore by providing voids which act as delamination initiation sites cracking can be provided between plies which allows greater flexibility towards the edge and therefore release of energy. These voids may incorporate uncured polymer matrix to act as a binder subsequent to delamination.

Abstract: Appropriate location of components such as vanes in gas turbine engines is important. These components are located between tangs held in interference engagement. There are four relatively rigid mounting points constituted by the tangs engaging respective opposed slots in a casing or a mounting ring. Only three point limitation with respect to displacement and rotation is adequate so provision of a fourth locator causes unnecessary constraint. A relatively resilient or compliant locator means it is possible to reduce the amount of constraint upon the component mounting whilst the other locators are adequate for positioning of the component.

Abstract: A blade arrangement 31 includes an array of radially extending blades 20, which may for example comprise a fan of a gas turbine engine for an aircraft. The blades 20 are mounted for rotation about a central axis X-X. The blade arrangement 31 further includes a damping arrangement 32 comprising means 34 for inducing an axi-symmetric magnetic field whose axis of symmetry coincides with the central axis X-X of rotation of the blades 20. The damping arrangement 32 is configured such that when the magnetic field is induced, any movement of the blade 20 other than pure rotation about the central axis results in the magnetic field causing a force to be exerted on the blade 20, the force resisting such movement. The damping arrangement may be provided with means for inducing the magnetic field only when there is an increased likelihood of vibration of the blades, for example when a foreign body has entered the air intake of the engine.

Abstract: A composite blade (26) comprises a root portion (36) and an aerofoil portion (38). The aerofoil portion (38) has a tip (42), a chord (C), a leading edge (44), a trailing edge (46), a suction surface (48) extending from the leading edge (44) to the trailing edge (46) and a pressure surface extending from the leading edge (44) to the trailing edge (46). The composite blade (26) comprises reinforcing fibres (52) in a matrix material (54). The aerofoil portion (38) adjacent the root portion (36) has regions (56, 58) at least at the leading edge (44) and trailing edge (46) comprising an asymmetric lay up of reinforcing fibres (52) and the aerofoil portion (38) adjacent the tip (42) having at least a region (62) at the mid-chord comprising an asymmetric lay up of reinforcing fibres (52).

Abstract: Previous testing of such test pieces as fan blades utilised in gas turbine engines has involved application of paint markings to identify fragments upon fragmentation of the test piece. Such fragmentation causes debris which will obscure paint markings and the paint itself may create clouds of dust obscuring visual images of the test piece under test conditions. By applying an image pattern comprising a number of lingering image patches, typically in the form of thermal patches upon the test components, and arranging the test components to be within an enclosure which is evacuated it will be appreciated that these image patterns linger and can be viewed by an appropriate monitor over a period at least of the test conditions. Utilisation of extra visual spectrum radiation images extends the period when clear images of the test piece under test conditions can be seen and analysed.

Abstract: With regard to hollow blades for turbine engines, it will be understood there is a problem with respect to percussive impact resulting in excessive distortion of the blade as well as potential failure as a result of blade tip bulging. By provision of ridges 107, 207, 307 which coincide and engage each other under impact, the extent of impact deformation is limited as well as a result of the narrowing between the ridges, a reduction in the possibility for fragmentary insert movement to bulge the cavity towards the tip 102, 202, 302 of a blade 100, 200, 300.

Abstract: A lightweight component in the form of an outlet guide vane 10 for a gas turbine engine. The vane 10 being filled with hollow metal ellipsoid members 12 in a solid metal outer 14.

Abstract: A platform arrangement for components such as aerofoil vanes 1, 21 is provided whereby the necessary rigidity provided by the platform does not create excess problems with respect to stress and strain fatigue failure at thinner sections of the component. The leading edge 6, 26 and trailing edge 7, 27 of a vane generally have a thinner section and so when subject to greater flex deformation such flex deformation may lead to premature failure through stress and strain. Structural configurations are now provided which comprise cut out windows 24 in the platform 22 beneath the edges 26, 27 or alternatively differential material types at these locations with a less stiff nature.

Abstract: A joint assembly 80 is provided for limiting an extension of the joint 80 in the direction of a load path derived from an impact comprising. In its simplest form, the joint 80 comprises a first member 40 having a portion 82 and a second member 50 having a portion 84, the portions 82, 84 overlapping one another and arranged generally parallel to one another and secured together via securing means 78. The securing means 78 is disposed through corresponding holes 92, 94 defined by portions 82 and 84. At least one of the overlapping portions 82 or 84 further defines, sequentially in the direction of extension a pocket 102. The securing means 78 comprises a deformable member 152, 178 and at least one of the overlapping portions 82, 84 comprises a wedge-shaped portion 74.

Abstract: A guide vane 100 is provided in which sheet portions 14, 15 are secured together to define an aerofoil profile 11. Between ends 20, 21 of the guide vane 100 there is a non linear variation in the maximum chordal thickness 13. Thus, greater maximum chordal thickness 13b in central portions of the guide vane 100 provide stiffness while ends 20, 21 which are generally formed from solid material have a smaller maximum chordal thickness such that a stiffer vane 100 can be provided with reduced material weight and therefore costs.

Abstract: A gas turbine engine rotor blade containment assembly comprises a generally cylindrical, or frustoconical, stiff containment casing (54), a generally cylindrical, or frustoconical, flexible structure (62) arranged within and spaced radially from the stiff containment casing (54) by crushable structures (64, 66) at axially spaced positions on the flexible structure (62). A viscoelastic material (68) is arranged to fill the space (70) between the stiff containment casing (54), the flexible structure (62) and the crushable structures (64, 66). The viscoelastic material (68) provides local stiffening of the blade containment assembly in the region of a fan blade (34) impact and provides energy dissipation by viscoelastic damping of the flexing of the flexible structure (62) and plastic deformation in the crushing of the crushable structures (64, 66).