Abstract: An aerofoil (39) for a gas turbine engine (10) includes one or more internal cooling passage (74,76) for receiving a flow of cooling fluid, and means (84,86,88,90) for inducing at least two vortices in cooling fluid flowing through each cooling passage. Fluid within the vortices has a radially outwards, screw-type motion. The two vortices produce effective cooling of the aerofoil, which may be an aerofoil of a turbine blade.

Abstract: A mold support (110) is disclosed for supporting a plurality of molds (109) in a casting apparatus having a first zone for introducing molten material into the molds and a second zone in which the material in the molds can cool. The mold support (110) is capable of relative movement from the first zone to the second zone thereby forming a solidification front in the material in each mold, whereby the solidification front moves through the material as the material solidifies. Each mold (109) defines a main axis (A) along which the solidification front can move and a centrally extending plane (P) perpendicular to the main axis. The mold support (110) comprises carrying means upon which the molds (109) can be disposed.

Abstract: A method of characterizing the vibration of a plurality of rotating blades (12), such as turbine blades in a gas turbine engine, includes the steps of providing a single strain gauge (16A) on a single one (12A) of the blades and providing either one or two tip timing probes (22A, 22B) on a casing surrounding the blades. The data from the strain gauge (16A) allows the vibration of the single blade (12A) to be fully characterized while the data from the tip timing probes (22A, 22B) allows the amplitudes or velocities of vibration of all the blades to be determined. The relationship between the data from the tip timing probes for the single blade (12A) and the stress data from the strain gauge may be established. This relationship may then be assumed to apply to all the blades, thus allowing the stresses induced in all the blades by the vibrations to be determined.

Abstract: The spiral heat exchanger (10) has a hot end (16) and a cold end (18) arranged at opposite ends of the heat exchanger (10). A first, relatively hot, fluid is supplied to the hot end (16) of the heat exchanger (10) and the first fluid is removed from the cold end (18) of the heat exchanger (10). The heat exchanger (10) comprises first and second metal sheets (12,14). Each metal sheet (12,14) has hot and cold edge portions (18,20) at the hot and cold ends (16,18) respectively of the heat exchanger (10). Both of the hot and cold edge portions (18) of the first metal sheet (12) are thinner and have a greater diameter than the remainder of the first metal sheet (12) such that both of the hot and cold edge portions (18) contact the hot and cold edge portions (20) of the second sheet (14). The hot and cold edge portions (18) of the first metal sheet are joined to the hot and cold edge portions (20) of the second metal sheet (14).

Abstract: A vane assembly 40 for use within a gas turbine engine has a main vane portion 42 with an internal cavity 44. A cavity insert 46 is located within the cavity 44, close to the wall 48 to define transpiration cooling paths. Cooling air leaves the insert 46 through apertures directed at the wall 48, to produce impingement cooling. The transpiration cooling paths are extended back to the trailing edge 66 by means of a fairing 54. The use of a fairing in addition to the insert allows more complicated cavity shapes to be filled.

Abstract: A rotor (38) for a gas turbine engine (10) has a low emissivity surface finish, or high emissivity surface finish (44), on at least a portion of its surface (42) to reduce temperature differentials between an upper portion and a lower portion of the rotor (38). This reduces bowing of the rotor (38) to allow the gas turbine engine (10) to be started without harmful vibrations of the rotor (38).

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) carrying a plurality of radially extending fan blades (34). A fan blade containment assembly (38) surrounds the fan blades (34) and the fan blade containment assembly (34) comprises a generally annular, or frustoconical, cross-section casing (52). At least one corrugated sheet metal ring surrounds the casing (52) wherein the corrugations extend with axial and/or circumferential components.

Abstract: An aerofoil (34) for a gas turbine engine is substantially solid, including a concave pressure surface (42) and a convex suction surface (40). The pressure surface (42) is provided with a projection in the form of an elongate fin (44) which extends in the radial direction of the aerofoil.

Abstract: A seal assembly for co-operation with a shaft comprising a seal member which, in use, in a first position, forms a seal against the shaft, a seal housing for retaining the seal member, means for applying a radially inward force to the seal member and means for applying a radially outward force to the seal member so that, when the means for applying a radially outwards force is operated, the seal member is dilated to a second position thereby enabling the seal assembly and the shaft to be fitted and disassembled in conventional manner, when the means for applying a radially outward force is removed the means for applying a radially inward force to the seal member returns the seal member to the first position.

Abstract: The present invention provides for an etch and mar resistant low VOC clear coating composition most suitable for use as a top clear coat in multi-layered OEM or refinish automotive coatings. The coating composition includes isocyanate, carbonate and melamine components. The isocyanate component includes an aliphatic polyisocyanate. The composition may be formulated as a two-pack or one-pack coating composition, wherein the isocyanate functionalities are blocked with a blocker such as a mono-alcohol.

Abstract: Cooling air enters a cooling path 56 from a supply at 58. The path follows a serpentine path consecutively through a number of components to be cooled, including guide vanes 38, 40, blades 30, a shroud 54 and a shaft 18. A valve 68 is provided at the downstream end of the path 56. The setting of the valve 56 modulates the flow and pressure of air along the path 56, without wholly preventing it.

Abstract: A brush seal (9) comprising a plurality of bristles packed together in a bristle layer (10). The bristles (10) extend from a first component (4) towards a facing surface of a second component (2). A movable plate (14) is disposed substantially parallel to and adjacent to the bristle layer (10). The plate (14) is movable relative to the bristle layer (10) in a direction parallel to the bristle layer (10). An edge (22) of the movable plate (14) adjacent to and facing the facing surface of the second component (2) is arranged to air ride on the facing surface of the second component (2). To promote the air riding the edge (22) of the movable plate (14) may be accordingly profiled. Alternatively the edge (14) may comprise an enlarged foot member (24) which defines an enlarged surface (26). Furthermore a recess (30) may be defined in the movable plate (14).

Abstract: A gas turbine engine nose cone (44) comprises a spinner (46) and a fairing (48). The spinner (46) has a conical upstream portion (50) and a cylindrical base portion (52). The fairing (48) is frustoconical and surrounds the base portion (52) of the spinner (46). The outer surface (68) of the fairing (48) forms a continuation of the outer surface (56) of the conical upstream portion (50) of the spinner (46). The outer surface (68) of the fairing (48) has a skin (80) to protect the fairing (48) from erosion and a protective member (86) extends around the upstream end (82) of the fairing (48) to retain the skin (80) on the upstream end (82) of the fairing (48).

Abstract: A thrust reverser cascade flow directing element (28) for a gas turbine engine (10) comprises a plurality of plates (30) and a plurality of flow directing vanes (40). At least one vane (40) extends between each pair of plates (30), the plurality of plates (30) are formed from sheet material and a single integral piece of ductile sheet material (32) forms the plurality of flow directing vanes (40). The single integral piece of ductile sheet material (32) has a plurality of longitudinally spaced apart apertures (36) and a plurality of longitudinally spaced apart sheet material portions (38) between the apertures (36) and the ductile sheet material (32) is bent at a plurality of longitudinally spaced positions such that each sheet material portion (38) defines one of the plurality of flow directing vanes (40).

Abstract: A turbine assembly (35) for a gas turbine engine (10) comprises a rotatable support arrangement (38) which comprises means for mounting thereon a plurality of turbine blades (36). The turbine assembly (35) defines flow path means (43) for a flow of cooling fluid therethrough. The flow path means (43) is connectable to a supply of relatively cold cooling fluid. The flow path means 43 is arranged such that the relatively cold cooling fluid is driven radially outwardly through the flow path means (43) substantially wholly by the centrifugal force generated the rotation of the turbine assembly (35) in operation. Relatively hot cooling fluid is displaced by the relatively cold cooling fluid radially inwardly through the flow path means (43).

Abstract: A particle vibration damper for a vibrating component is disclosed which comprises a body having a chamber and a plurality of particles, wherein the chamber is partially filled with the plurality of particles, the particle vibration damper, in use, is disposed on or in a vibrating component.

Abstract: A joint for joining a first metal sheet (10) to a second metal sheet (30) comprises a plurality of projections (18) on the edge (12) of the first metal sheet (10). The projections (18) are symmetrical about a plane extending perpendicularly to the plane of the first metal sheet (10). The symmetrical projections (18) are rotatable about an axis (X), the axis (X) is arranged in the plane of the first metal sheet (10) and in the plane extending perpendicularly to the first metal sheet (10). A plurality of slots (32) are provided through the second metal sheet (30). The symmetrical projections (18) extend through the slots (32) in the second metal sheet (30) and the symmetrical projections (18) are rotated such that the symmetrical projections (18) are arranged in the plane perpendicular to the plane of the first metal sheet (10) to clamp the second metal sheet (30) between the edge (12) of the first metal sheet (10) and symmetrical projections (18).

Abstract: An engine arrangement (10) comprises propulsion means for propelling the aircraft (12), and drive means to drive the propulsion means. The arrangement (10) further includes mounting means to mount at least the propulsion means in a location spaced above the upper surface of the aircraft (10). In the preferred embodiment, the propulsion means is in the form of tip driven fans (14), and the drive means comprises gas generators (34) mounted within the body of the aircraft (10) fixedly connected by intake ducts (36) to the underside of the aircraft (10).

Abstract: A control arrangement (34) for controlling the size of a gap (30) between first and second components. The control arrangement comprises first magnet means (38) to be provided on the first component, and second magnet means to be provided on the second component. The first and second magnet means are in magnetic interaction with each other across the gap (30) and the arrangement further includes control means (36) to control the size of the gap (30) the first magnet means (38) is a moveable in a first direction to cause the control means (36) to effect relative movement of the first and second components to increase the gap (30) and in a second direction to cause the control means (36) to effect relative movement of the first and second components to decrease the gap (30).

Abstract: A turbine blade for a gas turbine engine comprises an aerofoil having a suction and pressure side. The pressure side is provided with a reflex curvature at the aerofoil trailing edge region so as to reduce the thickness of the aerofoil in that region.