Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a blade containment portion (54) and a transition portion (58). The upstream portion (56) has a flange (52) connecting the metal containment casing (40) to an axially adjacent casing. The upstream portion (56) of the casing (40) has a greater diameter (D1) than the diameter (D2) of the blade containment portion (54) and the blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54). The transition portion (58) connects the blade containment portion (54) and the upstream portion (56) to transmit loads from the blade containment portion (54) to the upstream flange (42).

Abstract: A gas turbine engine casing includes a cavity 10 separated from the main flow 12 through the engine by an annular array of tip treatment bars 16. The tip treatment bars 16 are hollow which enables the bars to deform or break upon impact by a shed blade 4 or blade fragment. This allows the blade or blade fragment to pass through the tip treatment bars 16. Blade or blade fragments may then be contained within the cavity 10, or embedded in the casing 2, so preventing further damage to the engine. Entire blades, or large blade fragments, may be able to break through the tip treatment bars 16 and through the engine casing 2 so as to leave the engine entirely, causing the minimum of damage. The interiors of the tip treatment bars may be filled with a damping material, to minimize high cycle fatigue failure.

Abstract: A ducted fan gas turbine engine comprises a fan 12 mounted on a first shaft 16 and a compressor mounted on a second shaft 43. The shafts are coaxial both with each other and with said engine longitudinal axis 15. The shafts are mounted in roller bearings 21, 40. The first bearing 21 carrying the upstream part of the first shaft and the second bearing 40 carrying the upstream part of the second shaft. Both bearings are attached to a structure 26 which comprises an axial sleeve and a front frustoconical portion and a rear radially stiff but axially flexible frustoconical portion and is attached to the engine casing 36. The front frustoconical portion is frangibly attached to the front part of the axial sleeve 28. In the event of damage to the fan the frangible connection breaks and the first shaft is allowed to orbit about the engine longitudinal axis 15. The rear member 32 is swashingly non-linear flexible and accommodates the out of balance radial loads transferred to it as couples during such an incident.

Abstract: A locking device for use in retaining rotationally interengaged components in engagement with one another, the device comprising a body member (16) adapted to be located between adjacent interengaging formations (13, 14) on the components to prevent relative rotation of the components, and a wire retaining member (17) adapted to locate the body member in position, the retaining member comprising at least two hook-like members (18) engaged in spaced apertures (20) in the body member, each hook-like member having first and second portions (18B, 18C) extending beyond the body member into engagement with inner and outer surfaces of one of the components.

Abstract: A method of adding boron to a tungsten, or tantalum, containing titanium aluminide alloy to form a boride dispersion in the tungsten, or tantalum, containing titanium aluminide. A molten tungsten, or tantalum, containing titanium aluminide alloy is formed and tungsten, or tantalum, boride is added to the molten tungsten, or tantalum, containing titanium aluminide alloy to form a molten mixture. The molten mixture is cooled and solidified to form a tungsten, or tantalum, containing titanium aluminide alloy having a uniform dispersion of tungsten, or tantalum, boride particles substantially without the formation of clusters of tungsten, or tantalum, boride. The titanium aluminide alloy comprises between 0.5 at % and 2.0 at % boron.

Abstract: A magnetic clamping system and method for locating and holding a workpiece (6), in particular a non-magnetic workpiece (6) within a machine tool (2) and specifically a numerically laser cutting or drilling machine tool. The system comprising at least one magnet (9), and a fixture structure (7), attached to the machine tool, which cooperates and in use engages and supports a portion of the workpiece (6) mounted upon and within the fixture structure. At least one clamp member (11) made of a magnetic material is disposed opposite and facing the at least one magnet (9) with the magnet (9) so that the portion of the workpiece (6) is sandwiched between a clamp member (11) and the magnet (9). A magnetic attractive force between the clamp member (11) and magnet (9) providing a clamping force to hold the workpiece (6). The magnets (9) preferably permanent neodymium magnets.

Abstract: A cooling air flow control device for a gas turbine engine, the cooling air flow control device comprising a component, a cooling passage defined within the component and a shaped memory metal valve, the shaped memory metal valve disposed in the cooling passage to regulate, in use, the flow rate of a cooling air flow supplied, in operation, through the cooling passage wherein the shaped memory metal valve operates by changing shape to control the flow rate of the cooling air flow in response to the temperature of the component.

Abstract: A thrust reverser for a gas turbine engine (10) comprises cascade structures (20,22) mounted in an engine cowl (12). The cascades (20,22) comprise a plurality of air deflecting vanes (24) arranged in fixed space relationship. At least one of the cascades (22) translates between a first inoperative position where the cascade (22) is stowed radially inward of the cowl (12) and a second operative position where the cascade (22) is clear of the cowl (12) to expose the air deflecting vanes (24). When the trust reverser is operative blocker doors (26) pivot across an annular duct (13) defined between the core engine (11) and the cowl (12) to deflect the airflow passing therethrough to the exposed cascades (20,22) to produce a braking force.

Abstract: A gas turbine engine mounting arrangement (22) for attaching a gas turbine engine (10) to an aircraft via a pylon (18). The arrangement (22) comprising a main mounting means (23) which carries the engine loads under normal operating conditions, and an auxiliary mounting structure (25). The auxiliary mounting structure. (25) comprises a safety bracket (44), and auxiliary interconnection means (29) are arranged to connect the bracket (44) to the engine (10) independently of the main mounting means (23) and so that the auxiliary mounting structure (25) is substantially unloaded under normal operating conditions. The auxiliary interconnection means (29) and the safety bracket (44) engaging with the engine (10) and carrying substantial engine loads in the event of failure of the main mounting means (23). Preferably the safety bracket (44) is interposed and sandwiched between a main bracket (2), forming part of the main mounting structure (23), and the main bracket (2) and the pylon (18).

Abstract: A wall structure for a gas turbine combustor arranged to have a general direction of fluid flow therethrough includes inner and outer walls defining a space therebetween. The inner wall is made up of a plurality of tiles having axial edges aligned generally with the direction of fluid flow, a gap being defined between axial edges of adjacent tiles. Orifices are provided within the axial edges to direct leakage air passing through the gaps to give the leakage air a flow component in the general direction of fluid flow through the combustor.

Abstract: A method of manufacturing a gas turbine engine fan blade (10) comprises forming three metal workpieces (30,32,34). The metal workpieces (30,32,34) are assembled into a stack (36) so that the flat surfaces (38,42,46,48) are in mating abutment. Heat and pressure is applied across the thickness of the metal workpieces (30,32,34) to diffusion bond the metal workpieces (30,32,34) together to form an integral structure (80). The integral structure (80) is upset forged at one end (58) to produce an increase in thickness (82) for forming the blade root (26). The upset forged integral structure (80) is then hot creep formed and superplastically formed to produce the required aerofoil shape and the thickened end (82) is machined to form the blade root (26). The method enables thinner metallic workpieces with better microstructure to be used and increases the yield of metallic workpieces.

Abstract: A hollow liner (14) for fitting in the interior of a aerofoil vane (10) is crimped prior to fitting within the vane (10). Following fitting, the liner (14) is expanded so as to fit snugly within the vane (10). The technique permits the insertion of liners (14) into vanes (10) that are so configured as to preclude the insertion of conventional liners.

Abstract: A method for forming a body by deposition of a weld material comprises providing a welding head (12), and providing support means (34) to support the body. The welding head (12) and the support means (34) are connected to a supply of electricity to form an arc between the welding head (12) and the support means (34) to melt the weld material. The welding head (12) and the support means (34) are manipulated relative to each other to deposit the weld material and form a carrier member of the body. The carrier member is formed to carry a projecting member. The carrier member has a first portion (52) of a first predetermined thickness and a second portion (54) of a second predetermined thickness, the second predetermined thickness being greater than the first predetermined thickness. A projecting member is provided on the first portion.

Abstract: A method of forming a body by deposition of a weld material comprises providing a welding head (12), providing a foundation (52) upon which the body is to be formed. A weld material is supplied from a supply (22). The welding head (12) and the foundation (52) are connected to a supply of electricity to form an arc between the welding head (12) and the foundation (52) to melt the material. A raised member (74) is provided on the foundation (52). The method further includes manipulating the foundation (52) and the welding head (12) relative to each other to deposit the material on the raised member (74) and form the body.

Abstract: Apparatus for forming a workpiece by deposition welding comprises a welding head having at least two wires of weld material extending therethrough. Each wire is formed of weld material such that an electric current is passed therethrough, the wires become molten and form the weld material of the workpiece.

Abstract: A gamma titanium aluminide turbine blade (10) has a titanium oxide layer (20) on its outer surface. A protective coating (22) is applied to the aerofoil (12) and the platform (14) of the turbine blade (10) onto the titanium oxide layer (20). The protective coating (22) comprises a silicate glass having a chromium oxide filler. The protective coating (22) preferably comprise a boron titanate silicate glass having a chromium oxide filler. The oxide layer (20) adheres the protective coating (22) to the titanium aluminide turbine blade (10). The protective coating (22) provides oxidation and sulphidation resistance.

Abstract: An annular connector for connecting a rotor and shaft, has an axial cross sectional profile which approximates a serpentine form. The shape defines only obtuse angles and curves, so that during rotation of the assembly, stress variations in the connector are reduced, as are stress peaks.

Abstract: A nose cone assembly (30) for a gas turbine engine (10) comprising a spinner (34) having a generally conical forward portion (33) and a base portion (35). The base portion (35) of the spinner (34) is disposed radially inside an extended profile of the forward portion (33). Upon the base (35) there is a radially outwardly extending flange (41) through which mounting fasteners (40) are arranged to attach the spinner (34) to a fan hub (6) of the gas turbine engine (10). A fairing (36) is arranged to surround and cover the base portion (35) of the spinner (34) when the nose cone assembly (30) is fitted to the fan hub (6). The fairing (36) having a generally circumferentially and axially smooth continuous outer surface, the profile of which continues that of the forward portion 33 of the spinner 34. Fasteners (44) preferably extend from the fairing (36) in a substantially radially inwardly directed direction to attach the fairing (36).

Abstract: A tip treatment assembly is provided including a tip treatment bar (16) and retaining means (24, 26) which cooperate with a retaining element (30, 32) to retain the tip treatment bar (16) with respect to support means (44) of the assembly. The retaining means (24, 26) may comprise a hole through the bar (16) through which the retaining element, in the form of wire 30, 32, extends. The tip treatment bars (16) can be formed by stamping from a sheet of material such as a suitable alloy.

Abstract: A tip treatment bar component (16) for positioning within an annular cavity surrounding a fan (4) within a gas turbine engine comprises a longitudinal portion (30) with a platform (32, 34) at each end. The platforms (32, 34) of adjacent bars (16) abut each other laterally to provide a slot (28) between adjacent tip treatment bar components (16). A projection, for example a tang (24, 26), extends from the platforms (32, 34) in a direction away from the longitudinal portion (30) of the tip treatment component (16). This serves to locate the bar component (16) in the engine casing (2).