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 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 means 44 for abutment with 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: Appropriate location of components such as vanes (1, 21, 31, 41, 51, 61, 71, 81, 91) in gas turbine engines is important. These components are located between tangs (7, 27, 37, 47, 57, 67, 87, 97) held in interference engagement. There are four relatively rigid mounting points constituted by the tangs (7, 27, 37, 47, 57, 67, 87, 97) engaging respective opposed slots (8) in a casing (3) or a mounting ring (2). 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 (37, 47, 57, 67, 87, 97) means it is possible to reduce the amount of constraint upon the component mounting whilst the other locators (7, 27) are adequate for positioning of the component (1, 21, 31, 41, 51, 61, 71, 81, 91).

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 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 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 (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 fibres (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 gas turbine engine blade containment system (22) comprising a laminate material (24). The laminate material (24) comprising first and second layers (26, 28). The second layer (28) comprising a plurality of deformable members (30) extending transversely relative to the first layer (26), each deformable member (30) being encased in a crushable support material (34). The deformable members (30) comprising springs (32). The first layer (26) being an inner layer within the second layer (28). The first layer (26) comprising a hard material to blunt a broken blade. The springs (32) in the second layer (28) are compressed and the crushable support material (34) of the second layer (28) is crushed to absorb energy of a broken blade.

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: 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: 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: 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: 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: 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 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 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.