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 superconductor is placed inside a non-metallic cryostat filled with a cooling medium to maintain the superconductor in a superconductive state. A foil winding is connected in series with the superconductor by current leads and the cryostat is placed inside the winding. The winding has only a few turns in order to provide low inductance and low magnetic field when the superconductor is carrying its normal operating current. However in the event of a fault the increase in electrical current in the line winding causes a magnetic field to be generated parallel to the superconductor. The magnetic field generated exceeds the critical magnetic field of the superconductor. The magnetic field produced by the line winding triggers the superconductor to the resistive state.

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 and apparatus for creating dynamic shear within metallic and non-metallic test sample materials under controlled conditions related to high ballistic impact includes the use of test sample having an end face with non-symmetrically placed surfaces extending away from a leading edge of the end face.

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 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 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 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 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 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 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: A signal system 10 for a vessel such as an aircraft, incorporates nodes 16A, B, C arranged in a hierarchy. Low level nodes 16C are associated with plant 14 of the aircraft. Mid-level nodes 16B can communicate with groups of low level nodes 16C, grouped by proximity. A top level node 16A can communicate with all mid-level nodes 16B. Decision making capability is built into nodes 16, to allow decisions to be taken at the lowest hierarchical level at which all of the necessary information is available.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprising a generally cylindrical, or frustoconical, first metal casing (40) and a plurality of generally cylindrical, or frustoconical axially spaced second metal casing is (54, 56 & 58). The second metal casings (54, 56, 58) surrounded and abut the first metal casing (40). The first metal casing (40) is relatively hard and tough to contain a fan blade (34). The second metal casings (54, 56, 58 are relatively soft and lightweight to stiffen the first metal casing (40) and fan blade containment regions (A, B & C). The second metal casing (54, 56, 58) are arranged coaxially with the first metal casing (40). One of the second metal casing means (56) is in a plane (Y) containing the fan blades (34) of the gas turbine engine (10). The first metal casing (40) comprises steel and the second metal casing is (54, 56, 58) comprise aluminum. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: In a staged gas turbine combustion engine, each combustion chamber of an annular array of chambers includes at least a pilot and a main combustion stage each having an fuel injection nozzle; the method includes feeding fuel continuously in operation to the pilot stage nozzle and controlling the fuel flow to the main stage nozzle through a pulsing and dosing valve to control the fuel flow delivered to the main stage fuel nozzle.

Abstract: A gas turbine engine (10) is provided with a plurality of shafts (18, 19, 20) interconnecting its compressor portions (11, 12, 13) with its turbine portions (15, 16, 17). Each shaft (18, 19, 20) is supported in the hotter part of the engine by electromagnetic bearings (26, 28, 32) and by rolling element bearings (24, 27, 29, 31, 33) in the remainder of the engine (10). The arrangement provides a simplified liquid lubrication system for the rolling element bearings (24, 27, 29, 31, 33) which reduces the likelihood of lubricant thermal degradation of the lubricant.

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 (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The casing (40) comprises an annular hook (54) positioned axially upstream of the tip (37) of the fan blade (34). The annular hook (54) extends in a radially inwardly and axially downstream direction from the metal casing (40) towards the tip (37) of the fan blade (34). The metal casing (40) has a containment portion (A) downstream of the annular hook (54) and substantially in the plane of the leading edge of the tip (37) of the fan blade (34). The containment portion (A) of the casing (40) has a greater diameter than the diameter of the casing (40) at the annular hook (54) to restrain upstream movement of a detached fan blade (34).

Abstract: A nacelle assembly 40 adapted for mounting on a ducted fan gas turbine engine 20 comprising a generally annular body 46 having an air inlet 42 and an air outlet 44, the generally annular body 46 encircling a region of the engine 20 when working in operative association with the engine 20 and has a first attachment means 48 to a rigid member 58 and a second attachment means 52 to a casing assembly 34 on the engine 20 wherein the second attachment means 52 is frangible.