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).

Abstract: A metallic turbine blade (10) has a protective coating (22) applied to the shank (16) and the root (18) of the turbine blade (10). The protective coating (22) comprises a chromized coating (24) diffused into the surface of the metallic article (10) and a glass coating (26) on the chromized coating (24). The glass coating (26) comprises a silicate glass, preferably having a chromium oxide filler. The glass coating (26) preferably comprises a boron titanate silicate glass having a chromium oxide filler. The protective coating (22) provides oxidation and sulphidation resistance for the shank (16) and root (18) of the turbine blade (10).

Abstract: An aerofoil (22,24), preferably of a high lift, highly loaded design, for an axial flow turbo machine (10). The aerofoil having a span, a leading edge (LE), a trailing edge (TE) and a cambered sectional profile comprising a pressure surface (30,72) and a suction surface (28,70) extending between the leading edge (LE) and trailing edge (TE). The aerofoil (22,24) having at least one aerofoil cross bleed passage (36,37,78,80) defined in the aerofoil (22,24) which extends from the pressure surface (30,72) through the aerofoil (22,24) to the suction surface (28,70). The at least one passage (36,37,78,80) preferably disposed generally at a location on the suction surface (28,70) at which boundary layer separation from the suction surface (28,70) would normally occur.

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 method for analysing a thermal paint includes the step of defining a color space in at least two dimensions, the dimensions representing different colors and/or luminance values. The location of reference points, representing calibration data, is defined within the color space for comparison with color information relating to a part to be analysed. Each pixel of an image of the part to be analysed is given a location in color space and the nearest calibration reference point is determined. In this way, the approximate temperature of that point can be established.

Abstract: A vibration damper comprises a bearing located between two rotatable shafts and an axially extending oil entry surface positioned radially inwardly of the radial outer surface of the bearing outer race and axially adjacent one end of the bearing. The outer surface of the bearing outer race forms a clearance between itself and the bearing housing for the ingress of a film of oil directed from the oil entry surface. A weir is positioned at the axially opposite end of said bearing and radially inward of the bearing outer race and oil entry surface such that in use the radial height difference between oil entry surface and weir provides a pumping force and supplies a film of oil between the bearing outer race and its casing.

Abstract: A method of manufacturing an article (10) from two metal workpieces (30,32), the first and second metal workpieces (30,32) define the outer profile of the article (10). The second workpiece (32) is thicker than the first workpiece (30). The first workpiece (30) has two flat surfaces (34,36) and the second workpiece has one flat surface (38) and one machined (42) surface (40). Stop of is applied to surface (34) to prevent diffusion bonding at preselected areas. The workpieces (30,32) are assembled into a stack (58) so that the flat surfaces (34,38) are in abutment. Heat and pressure are applied to diffusion bond the workpieces (30,32) together to form an integral structure (60). The integral structure (60) is placed in a die such that the first workpiece (30) faces a convex surface of the die. The integral structure (60) is heated to cause the integral structure (60) to be hot creep formed on the convex surface of the die.

Abstract: The fire resistance of components normally subject to a standard flame test is evaluated using low temperature model tests. Reynolds number, Froude number, the pressure ratio of the flame to the surrounding atmosphere are matched in the modelling parameters and account is taken of the Prandtl number. Also, the model flow is controlled so as to match the change of momentum flux ratio in the flow through the burner used in a standard flame test. The method allows testing of easily produced model components at temperatures less than 90° C., with savings of cost and time.

Abstract: To reduce profile losses in the aerofoils of low pressure turbines of axial flow reaction turbines, the wake-passing effects on a row of aerofoils downstream of a first row of aerofoils are enhanced by providing a region of roughness on part of the suction surface of each aerofoil of the downstream row. The region of roughness has its leading edge between the location of the geometric throat on the suction surface and a location 75% of the surface perimeter from the leading edge. The region extends over at least 3% of the suction surface perimeter. The use of such a region of roughness is particularly effective in improving the performance of very high-lift aerofoils (coefficient of lift of 1.1 or greater).

Abstract: A three stage lean burn combustion chamber (28) comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (76,78,80,92). The secondary fuel and air mixing duct (80) has passages (80A) and apertures (90A) at its downstream end to supply air and fuel into the secondary combustion zone (40) at a first position in the at least one combustion zone (40) and the secondary fuel and air mixing duct (80) has passages (80B) and apertures (90B) at its downstream end to supply air and fuel into the secondary combustion zone (40) at a second position in the secondary combustion zone (40) downstream from the first position. This axial distribution of fuel in the combustion zone (40) reduces the generation of harmful vibrations in the combustion chamber (28).

Abstract: A method and apparatus for anticipating the onset of or measurement of a deleterious condition in a bearing comprising monitoring the bearing for electrostatic activity indicative of the onset of the condition using one or more electrostatic sensors, processing signals derived from the sensors and employing same to activate alarm means, alter bearing loading to reduce or eliminate the source condition or shut down the equipment to prevent further bearing damage.

Abstract: The shroud (26) of a shrouded turbine blade (22) has a convex leading edge (30) the apex of which extends towards the trailing edge of the shroud (24) of an adjacent, preceding guide vane. The extension enables the containment of a reverse flow of high pressure gas from the juncture of the blade aerofoil and shroud, until the pressure has dropped sufficiently to avoid nullifying gas flow from the corresponding portion of the guide vane (20) and thereby reduces excessive hot gas leakage from between the opposing shrouds edges.

Abstract: A method of performing an operation comprises operation steps and evaluation steps.

Abstract: A wall element (29A, 29B) for a combustor (20) of a gas turbine engine (10). The wall element (29A, 29B) defines an axis. In use, the axis is arranged generally parallel to the principal axis of the engine (10). In one aspect, the length of the wall element (29B) along the axis is at least substantially 20% of the length of the wall element (29B) transverse to the axis. In another aspect, the wall element (29A, 29B) has a first pair of opposite edges extending transverse to the axis and a second pair of opposite edges (48, 50) extending transverse to the first pair, at least one of the second pair of edges (48, 50) being angled relative to the axis of the wall element (29A, 29B).

Abstract: A gas turbine engine casing includes a tip treatment ring 20 which comprises end supports 18 which support tip treatment bars 16. Each tip treatment bar 16 is mounted at its ends within the respective end supports 18 by damping elements 22, which isolate the bar 16 from the end supports 18. The elements 22 damp vibrations induced in the tip treatment bars 16 so inhibiting the initiation of high cycle fatigue cracking.

Abstract: A nickel based single crystal superalloy comprising 6-11 wt % cobalt, 4.7-5.7 wt % chromium, 2.4-3.0 wt % molybdenum, 3.0-3.8 wt % tungsten, 3.0-3.8 wt % rhenium, 5.5-7.0 wt % aluminium, 5.0-6.0 wt % tantalum, 0.5-1.0 wt % niobium, 0-0.2 wt % hafnium, 0-150 ppm carbon, 0-100 ppm yttrium, 0-100 ppm lanthanum, 0-5 ppm sulphur and the balance nickel plus incidental impurities. The nickel based single crystal superalloy is suitable for use as a gas turbine engine turbine blade or turbine vane. It is of particular use on cooled turbine blades and turbine vanes which have ceramic thermal barrier coatings, because the superalloy is compatible with the ceramic thermal barrier coating to minimize spalling. The superalloy has lower density than other second generation single crystal superalloys but similar creep strength and oxidation resistance.

Abstract: A method of controlling bearing load comprises sensing the approach of a no-load condition in a bearing, generating a single in response thereto and applying a load to the bearing in a direction opposite to the direction of loading before the approach of said no-load condition whereby to take the bearing through the no-load condition in a rapid and controlled manner to a condition in which it is loaded in the opposite direction. Preferably the load is applied to the bearing by an electrostatic actuator in the form of a separate electromagnetic bearing surrounding the shaft.

Abstract: In one aspect the invention provided a friction welded component comprising at least two structural support members (2, 6) spaced apart and connected by a plurality of first (4) and second (12) studs. Each of the first studs are friction welded at one end to a first of the support members (2) and at the opposite end thereof to one end of a respective second stud. Each of the second studs are friction welded to a respective first stud and a second of the support members along a continuous annular friction welded joint (8, 10, 14) having a frusto-conical configuration.

Abstract: A nickel alloy composition, having particular utility for forming an oxidation resistant blade tip on a turbine blade, preferably has a composition substantially as defined by the nominal composition Cr 4.5 wt %; Al 6 wt %; Co 4 wt %; Ta 6 wt %; Re 4 wt %; Hf 0.15 wt %; C 0.05 wt %; Si 0.1 wt %.; B 0.005 wt %; W 2 wt %; La 0.003-0.005 wt %; and Y 0.003 to 0.005 wt %; the remainder being nickel.