Abstract: Where gas turbine engine structure eg combustion equipment, is to be air impingement cooled, the surface which receives the air jets is so shaped as to produce boundary layer separation zones 34, 38 and 44 in the cooling air, as it spreads across the surface. Mixing of the boundary layer with the remainder of the air flow results, followed by the re-establishment of the boundary layer. The new boundary layer is cooler than the original layer and so provides more effective cooling.

Abstract: A method of prestressing a component (30) includes the use of an electrical discharge or current to produce a plasma (39) within a medium (32) located adjacent the component (30). The plasma generates a shock wave which impacts a surface of the component to produce a region of compressive residual stress within the component.

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: An irreversible temperature indicating paint comprise 15 to 18 wt % cobalt ammonium phosphate, 5.0 to 6.0 wt % sodium alumino sulpho silicate, 8.5 to 10.5 wt % silica, 7.0 to 8.5 wt % alumina, 0.5 to 0.7 wt % toluidine red, 35 to 43 wt % acrylic resin and 19 to 23 wt % silicone resin excluding solvent. A particular irreversible temperature indicating paint comprising 16.4 wt % cobalt ammonium phosphate, 5.5 wt % sodium alumino sulpho silicate, 9.4 wt % silica, 7.8 wt % alumina, 0.6 wt % toluidine red, 39.1 wt % acrylic resin and 21.1 wt % silicone resin excluding solvent. The irreversible temperature indicating paint has at least six color changes in the temperature range 550° C. to 1250° C. The irreversible temperature indicating paint is used to determine the temperatures to which various parts of turbine blades, turbine vanes or other components are subjected in operation of a gas turbine engine.

Abstract: An electrical conductor winding (70) comprises a plurality of laminates of electrical insulation (72). Each laminate of electrical insulation (72) has a slot (78) on one surface (76). Each laminate of electrical insulation (70) has an electrical conductor (82) arranged in the slot (78). An electrical connector (84) connects the electrical conductor (82) in one laminate of electrical insulation (78) with the electrical conductor (82) in an adjacent laminate of electrical insulation (72). The laminates of electrical insulation (72) are arranged such that the surface (74) of one laminate of electrical insulation (72) abuts and is bonded to the surface (76) of an adjacent laminate of electrical insulation (12). The laminates of electrical insulation (12) comprise a glass-ceramic material. The electrical conductor windings allow active magnetic bearings, electric motors and electric generators to be used at temperatures up to 500° C. for example in gas turbine engines.

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: In a method for machining a workpiece, and for identifying a sequence of steps for machining a workpiece, a final shape required for the whole or part of the workpiece (is identified at 34), a three dimensional envelope shape is generated around the required shape (36), differences between the current shape and the envelope shape are generated at (38) to allow a sequence of cuts to be generated at (40) and a tool path at (42) from which the current stock part can be cut down to the envelope shape. The definition of the current stock is then updated at (44) and the sequence repeated until the workpiece has been machined to the required shape.

Abstract: A device and method for monitoring variations in the distance between an object and a datum, comprising: a light source, an astigmatic system for projecting an astigmatic image of the light source onto the object, the astigmatic system including the datum; and image shape detector means for detecting changes of shape of the astigmatic image on the object due to the variation in the distance between the object and the datum, and for producing a monitor signal whose value is dependant on the shape of the astigmatic image on the object: the astigmatic system being arranged such that light reflected from the astigmatic image on the object passes back through the astigmatic system and is thereby projected onto a light receiving surface of the image shape detection means as an astigmatic image of the astigmatic image on the object; wherein the astigmatic system comprises a zone plate.

Abstract: An irreversible temperature indicating paint comprises 25 wt % to 50 wt % cobalt silicate, 0 wt % to 20 wt % alumino silicate, 0.5 wt % to 5 wt % toluidine red, 25 wt % to 40 wt % acrylic resin and 10 wt % to 20 wt % silicone resin excluding solvent. The solvent comprises a mixture of 80% 1-methoxy-2-propanol and 20% dipropylene glycol monomethyl ether. The irreversible temperature indicating paint has at least five colour changes in the temperature range 1050° C. to 1350° C. A particular irreversible temperature indicating paint comprises 33.8 wt % cobalt silicate, 16.9 wt % alumino silicate, 2 wt % toluidine red, 30.8 wt % acrylic resin and 16.5 wt % silicone resin excluding solvent. The irreversible temperature indicating paint is used to determine the temperatures to which various parts of turbine blades, turbine vanes or other components are subjected in operation of the gas turbine engine.

Abstract: A gas turbine engine fan blade (26) comprises a root portion (40) and an aerofoil portion (42). The aerofoil portion (42) has a leading edge (44), a trailing edge (46), a concave metal wall portion (50) extending from the leading edge (44) to the trailing edge (46) and a convex metal wall portion (52) extending from the leading edge (44) to the trailing edge (46). The aerofoil portion (42) has a hollow interior (54) and the interior (54) of the aerofoil portion (42) is at least partially filled with a vibration damping material (56). The vibration damping material (56) comprises a material having viscoelasticity for example one formed by mixing an amine terminated polymer and bisphenol a-epichlorohydrin epoxy resin.

Abstract: A wall element (29) for a wall structure (21) of a gas turbine engine combustor (15). The wall element (29) comprises a main member (36) with an upstream edge region (30) and a downstream edge region (31). A plurality of heat removal members (38) are provided on the main member (36). The downstream edge (35) of the wall element and/or the downstream facing surface of the heat removal members closest to the downstream edge (35) are provided with a thermally resistant coating.

Abstract: A method of manufacturing a ceramic matrix composite comprises forming a slurry comprising a ceramic sol, filler particles and a solvent and forming laminates of fibers (12). The laminates of fibers (12) are impregnated with the slurry and are stacked (14) on a mold (10). The stack (14) of laminates of fibers (12) is covered by a porous membrane (16), a breather fabric (18) and a vacuum bag (20). The vacuum bag (20) is evacuated and is heated to a temperature of 60° C. for 10 hours to produce a ceramic matrix composite. The ceramic matrix composite is then heated to a temperature of 1200° C. at atmospheric pressure to sinter the ceramic matrix composite.

Abstract: A gas turbine engine (10) comprises a fan rotor (26) having a plurality of fan blades (24) and an apparatus (34) for detecting damage to the fan blades (24). The apparatus (34) comprises a transducer (36) arranged to detect the pressure in the gas flow around the fan blades (34) and to produce a pressure signal. A speed sensor (48) is arranged to measure the speed of rotation of the fan rotor (26) and to produce a speed signal. A processor unit (40) analyses the pressure signal and the speed signal to detect changes in the amplitude of the pressure signal, which occur at multiples of the rotational frequency of the fan rotor (24). The changes indicate differences in pressure between the gas flow around a damaged fan blade (24) and the gas flow around the remainder of the fan blades (24). The processor unit (40) sends a signal indicative of damage to a fan blade (24), if the difference in pressure is above the predetermined level, to an indicator device (44, 46).

Abstract: A vibration damper 46 for a gas turbine engine has convergent friction surfaces 48a and 48b and is located radially inward of the platforms 20a and 20b of two adjacent turbine blades. The angle subtended by the friction surfaces 48a and 48b is smaller than that subtended by the angled faces 22a and 22b associated with the platforms 20a and 20b. The center of mass of the damper 46 lies in a plane bisecting the angle subtended by the friction surfaces 48a and 48b. In use the damper 46 is urged radially outwards by centrifugal force so that at least one of the friction surfaces 48a and 48b makes planar contact with at least one of the angled faces 22a and 22b. Vibrational energy is dissipated by the resultant sliding movement between the friction surfaces 48a and 48b and the angled faces 22a and 22b. A secondary vibration damping mechanism arises from the oscillation of the damper 46 between the platforms 20a and 20b.

Abstract: An overthrust protection system for a gas turbine in which fuel is supplied at a controlled flow rate to a combustor (4) of the gas turbine. The overthrust protection system uses a bypass valve (24) to divert fuel flow away from the combustor (4). The opening of this valve is controlled by an overthrust controller (28) to divert fuel flow in response to a detected overthrust condition of the gas turbine. Diverting some of the fuel flow in this manner brings about a reduction in the thrust, allowing the overthrust condition to be controlled without resorting to a complete shut down of the gas turbine.

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 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 hot forming die (10) has forming surfaces (16, 18) and the forming surfaces have a nickel oxide layer (20). A stop off material coating (22) builds up on the nickel oxide layer (20) in operation. A method of cleaning the forming surfaces (16, 18) of the hot forming die (10) comprises washing the forming surfaces (16, 18) of the hot forming die (10) in water to remove the stop off material (22) without removing the nickel oxide layer (20) from the forming surfaces (16, 18). The forming surfaces (16,18) may be soaked in water (26) and/or a jet (30) of pressurised water may be directed onto the forming surfaces (16, 18). The stop off material is quickly removed without damaging the forming surfaces (16, 18). The retention of the nickel oxide layer (20) improves the quality of the hot-formed articles and the interval between cleaning of the hot forming die (10) is increased.

Abstract: A sealing element (30) for a turbine of a gas turbine engine includes a radially inner surface region provided with an integrally formed seal structure (38) comprising a plurality of radially inwardly projecting walls (40). The walls may be abradable and may define cells (44) for receiving an abradable sealing material.

Abstract: A stage of turbine blades (40) in a gas turbine engine (10) is surrounded by an array of shroud segments (42). The upstream ends of the segments (42) have plenum chambers (54) into which cooling air is fed from a compressor (12) via one hole (66) of a pair of holes, the other being numbered (68). Air from the plenum chambers (54) passes out to film cool the interior surface of each respective segment (42). Air from holes (68) passes out to convection cool the exterior surface of each segment (42), which effect is enhanced by the provision of ribs (80) and fences (82).