Abstract: A method of forming a blade or vane for a gas turbine engine comprising: attaching a first outer layer to a first two-dimensional array of attachment areas on a first surface of an intermediate layer; attaching a second outer layer to a second two-dimensional array of attachment areas on a second surface of the intermediate layer opposite the first surface; and increasing a separation between at least a portion of the first and second outer layer to thereby deform the intermediate layer into a corrugated structure having corrugations in first and second directions.

Abstract: A vane assembly for a gas turbine engine comprises at least one vane, a first platform and a second platform. The at least one vane has a first end and an opposite second end with the first end of the at least one vane being secured in a corresponding first aperture in the first platform, and the second end of the at least one vane being secured in a corresponding second aperture in the second platform. The at least one vane intersects the first platform at an acute angle, and the at least one vane intersects the second platform at an obtuse angle with the acute angle and the obtuse angle being supplementary to one another.

Abstract: A rig for testing an aerofoil component in four-point bending, which includes first and second fixtures movable relative to each other to apply a cyclically varying load to opposing aerofoil surfaces of the component. The first fixture has two first loading formations positioned to apply the load to one surface of the component, the formations being spaced apart in the spanwise direction of the component. The second fixture has two second loading formations positioned to apply the load to the other surface of the component, which are also spaced apart in the spanwise direction of the component and positioned relative to the first loading formations to apply the load in a four-point bending arrangement. The second fixture is continuously adjustable so that its loading formations move relative to each other in the direction of application of the load to balance the load between the formations during the cyclical load variation.

Abstract: A rig for testing an aerofoil component in four-point bending, which includes first and second fixtures movable relative to each other to apply a cyclically varying load to opposing aerofoil surfaces of the component. The first fixture has two first loading formations positioned to apply the load to one surface of the component, the formations being spaced apart in the spanwise direction of the component. The second fixture has two second loading formations positioned to apply the load to the other surface of the component, which are also spaced apart in the spanwise direction of the component and positioned relative to the first loading formations to apply the load in a four-point bending arrangement. The second fixture is continuously adjustable so that its loading formations move relative to each other in the direction of application of the load to balance the load between the formations during the cyclical load variation.

Abstract: A vane assembly for a gas turbine engine comprises at least one vane, a first platform and a second platform. The at least one vane has a first end and an opposite second end with the first end of the at least one vane being secured in a corresponding first aperture in the first platform, and the second end of the at least one vane being secured in a corresponding second aperture in the second platform. The at least one vane intersects the first platform at an acute angle, and the at least one vane intersects the second platform at an obtuse angle with the acute angle and the obtuse angle being supplementary to one another.

Abstract: A method of manufacturing a composite rotor disc (2) in which, following diffusion bonding of a ceramic core section (4) of the rotor disc (2) to a metallic outer section (6) of the rotor disc (2) at an elevated temperature, the rotor disc (2) is allowed to cool to a predetermined temperature and is held at this predetermined temperature for a predetermined period of time. While the rotor disc (2) is held at this predetermined temperature it is loaded in a radially outward direction. The rotor disc (2) is then allowed to cool to a temperature below the predetermined temperature.

Abstract: A method of manufacturing a composite rotor disc (2) in which, following diffusion bonding of a ceramic core section (4) of the rotor disc (2) to a metallic outer section (6) of the rotor disc (2) at an elevated temperature, the rotor disc (2) is allowed to cool to a predetermined temperature and is held at this predetermined temperature for a predetermined period of time. Whilst the rotor disc (2) is held at this predetermined temperature it is loaded in a radially outward direction. The rotor disc (2) is then allowed to cool to a temperature below the predetermined temperature.

Abstract: Bracing means for supporting a moveably mounted or flexible component during a machining operation that has an expandable member adapted to fit between a component and a support surface. When the expandable member is expanded a bracing member will resist movement and/or flexing of the component caused by forces arising during the machining operation. The expandable member is deformable to substantially conform to the shape of the component and the support surface.

Abstract: Bracing means for supporting a moveably mounted or flexible component during a machining operation that has an expandable member adapted to fit between a component and a support surface. When the expandable member is expanded a bracing member will resist movement and/or flexing of the component caused by forces arising during the machining operation. The expandable member is deformable to substantially conform to the shape of the component and the support surface.

Abstract: Bracing means for supporting a moveably mounted or flexible component during a machining operation that has an expandable member adapted to fit between a component and a support surface. When the expandable member is expanded a bracing member will resist movement and/or flexing of the component caused by forces arising during the machining operation. The expandable member is deformable to substantially conform to the shape of the component and the support surface.

Abstract: A frangible coupling for interconnecting parts comprising a first ring and a second ring coaxially arranged relative to each other and axially joined via an annular array of fuse ligaments equidistantly spaced apart from each other. The ligaments are configured to fail when an abnormal radial load of a predetermined value causes the first and second ring to move out of their coaxial relationship. When all of the fuse ligaments are severed, the communication between the first and second rings is severed. This allows the first ring to move independently of the second ring, preventing the out of balance load on the first ring being communicated to the second ring.

Abstract: A tip treatment bar (16) for use in a gas turbine engine casing is provided with a cavity which is open on one side of the bar (16), the cavity being filled with a damping material (24). The bars (16) may be open ended. The damping material (24) reduces vibration amplitudes in the bar (16) caused by aerodynamic excitation, thus reducing high cycle fatigue of the tip treatment bars (16).

Abstract: A frangible coupling for interconnecting parts comprising a first ring and a second ring coaxially arranged relative to each other and axially joined via an annular array of fuse ligaments equidistantly spaced apart from each other. The ligaments are configured to fail when an abnormal radial load of a predetermined value causes the first and second ring to move out of their coaxial relationship. When all of the fuse ligaments are severed, the communication between the first and second rings is severed. This allows the first ring to move independently of the second ring, preventing the out of balance load on the first ring being communicated to the second ring.

Abstract: A tip treatment bar (16) for a gas turbine engine has a damping coating (24) applied to one or more sides. The coating (24) may be a hard ceramic, for example Magnesia Alumina Spinel, and this may be plasma sprayed directly on to the bar (16) or applied to a substrate (26) which may then be applied to the bar (16).

Abstract: A tip treatment bar (16) for use in a gas turbine engine casing is provided with a cavity which is open on one side of the bar (16), the cavity being filled with a damping material (24). The bars (16) may be open ended. The damping material (24) reduces vibration amplitudes in the bar (16) caused by aerodynamic excitation, thus reducing high cycle fatigue of the tip treatment bars (16).

Abstract: A tip treatment bar (16) for a gas turbine engine has a damping coating (24) applied to one or more sides. The coating (24) may be a hard ceramic, for example Magnesia Alumina Spinel, and this may be plasma sprayed directly on to the bar (16) or applied to a substrate (26) which may then be applied to the bar (16).