Abstract: A plurality of workpieces are assembled into a stack. Two of the workpieces are machined on one surface at predetermined positions from the edges of the workpieces. The workpieces are stacked so that the machined surfaces of the workpieces face away from each other. Two workpieces are provided with a stop off material on one of their surfaces to prevent diffusion bonding at predetermined positions. The edges of the workpieces are welded together and a pipe is welded to the stack to interconnect with the stop off material to form a sealed assembly. The workpieces are diffusion bonded together. The integral structure formed is heated and opposite ends of the integral structure are twisted relatively to contour the integral structure to a predetermined shape. The integral structure is heated and the machined portions on the workpieces are isothermally forged and then the twisted integral structure is internally pressurized to superplastically form one workpiece to produce an article, e.g. a fan blade.

Abstract: A coating composition comprising an alloy having the formula RCrAlR'R" wherein R is nickel, cobalt or the like; R' is yttrium or hafnium and R" is tantalum, rhenium and/or platinum, preferably mixed with an oxide dispersion such as alumina, to provide an improved class of coatings suitable for operating in high temperature oxidizing environments.

Abstract: A fan blade assembly for a ducted fan gas turbine engine comprises a hub which carries an annular array of radially extending fan blades. Each fan blade has a root portion which locates in a corresponding groove in the hub. Aligned radial slots in the root portion and groove accommodate a U-shaped key which provides axial retention of the root portion in its groove. A flat spring attached to the underside of the root portion maintains the key in position and permits manual deflection of the key to assist blade assembly and disassembly.

Abstract: A shaft (12) is connected to a disc (20) by a coupling which comprises a spigot (14) which extends from the shaft (12) and locates in a recess (22) in the disc (20). At ambient temperature there is an interference fit between the spigot (14) and the recess (22) which maintains alignment of the disc (20) and the shaft (12) so that they rotate concentrically. At elevated temperatures the interference fit loosens and alignment of the disc (20) and the shaft (12) is maintained by a plurality of ridges (27) on the face (25) of a flange (24) on the disc (20). The ridges (27) have a large radial component and engage with the confronting face (17) of a flange (16) on the shaft (12) to prevent lateral movement between the shaft (12) and the disc (20).

Abstract: A heat pipe, which may for instance be in the form of an aerofoil blade for a gas turbine engine, is provided with capillary means which is in two different forms. One form is constituted by a porous material and the other constituted by grooves formed in the surface of the internal wall of the heat pipe.

Abstract: A gas turbine engine heat exchange apparatus comprises a first heat exchanger (12), which is arranged to receive oil from a lubricating oil system of the engine and put the oil in heat exchange with either a cooling or heating medium (17, 21), and a second heat exchanger (26), which receives the oil from the first heat exchanger, puts the oil in heat exchange with fuel in the engine's fuel supply system and discharges the oil back to the lubricating oil system. A valve or valves (20, 24) are used to control flow from high and low temperature sources of the heat exchange medium through the first heat exchanger (12) such that the direction of heat flow between the oil and the cooling or heating medium can be changed as necessary to prevent the fuel temperature from straying beyond predetermined limits.

Abstract: In the event of incipient engine speed or loss of normal engine control fuel flow rate in a jet engine is limited by a speed governor. A speed governor arrangement with optionally selectable emergency speed control is described. Maximum engine speed limits are stored in ROM and compared with sensed values. When an overspeed threshold is reached an output is produced to reduce fuel flow rate. If normal engine control is lost an alternative governor mode can be selected in which a second fuel flow rate is read from ROM as determined by the pilot's speed lever setting. A lowest wins circuit selects the lower of this emergency control or the governor limit signal to drive the fuel metering unit.

Abstract: A gas turbine engine combustor 20 has a wall structure 22 including an outer wall 24 having a plurality of wall elements 26 attached thereto. Each wall element 26 has a flange 27 around its periphery which defines a chamber 28 between each wall element and the outer wall 24. Holes 30 in the outer wall 24 permit the flow of cooing air into each chamber 28 to provide impingement cooling of the wall elements 26. Holes 30 in the wall elements 26 permit the exhaustion of cooling air from the chambers to provide film cooling of the wall elements 26.

Abstract: A rake for measuring pressures or temperatures in a gas flow includes an elongate member which has a single recess in the front face thereof. Within the recess are located a plurality of sensors. The recess defines a single cavity around the sensors and bleed slots are provided in the side faces of the member which are in fluid communication with the cavity. The sensors are connected to a remote pressure or temperature measuring system. In operation the rake is supported in a gas flow a proportion of which passes into the cavity. The cavity decelerates the gas and removes any swirl so that the sensors take an accurate measurement of the pressure or temperature. The gas then exits the cavity through the bleed slots.

Abstract: In a mounting for a gas turbine engine, the engine is pivotally mounted about its center of gravity to an aircraft support structure by means of a pair of radially outwardly extending trunnions. Each trunnion is connected to a central conical section of engine casing by means of an axially extending intermediate support member. The support member is secured to the downstream end of the conical casing section at an axial location offset from the trunnion. During operation distortion of the less stiff conical section is reduced by virtue of the resultant shear force distribution across the casing.

Abstract: A feed arrangement for a rotating bar includes a rotary drive carrying a series of axially spaced clamping devices comprising collets through which the bar passes. The clamping devices are axially reciprocated in phased sequence and the collet of each device closes upon the bar while the device is moving forwards. The sequence is so controlled that at all times the bar is clamped by the collet of at least one forwardly moving device and is thereby continuously fed forwards while being rotated by the active clamping device.

Abstract: This invention provides a valve for diverting fluid flow in a tandem fan type engine. The valve comprises an axially displaceable sleeve 38 which is provided with two circumferentially extending sealing lands 40, 42 which in operation cooperate with two locating surfaces 50, 52 provided on the duct 22 between the first and second compressors 20, 24 of said engine to seal an outlet opening 36. The free unsprung length between the sealing lands 40, 42 is greater than the length between the locating surfaces 50, 52. Gas pressure is used to compress a contractable portion 54 in said sleeve and hence ensure sealing takes place between both pairs of locating surfaces and sealing lands. This apparatus ensures centralisation of the upstream and downstream ends of the sleeve 38 is independently obtainable and accommodates thermal expansion differentials between the sleeve 38 and the duct 22.

Abstract: A turbine nozzle assembly includes an annular array of nozzle guide vanes located downstream of a combustor discharge casing. Each nozzle guide vane includes an aerofoil portion which is cast integrally with a radially inner platform and a radially outer platform. The radially outer platform of each nozzle guide vane has an extension to provide a smooth transition of the gases from the combustor discharge casing to the nozzle guide vanes. Two rows of cooling holes are provided in the extension to film cool the inner surface of the platform. A method is described to calculate the diameter of each of the cooling holes so that a uniform flow of cooling air passes over the inner surface of the each platform.

Abstract: A gas turbine engine fuel injection apparatus comprises a fuel spray atomizer which directs a fuel spray onto the radially inner surface of an annular flow deflector. The fuel flows in a film over the flow deflector surface towards an annular lip at the downstream end of the deflector. Swirling air flows are directed over the radially inner and outer surfaces of the flow deflector so as to atomize the fuel as it leaves the annular lip. The fuel is evaporated in the swirling airflows and thoroughtly mixed with the airflows in a mixing duct before being discharged into a combustion chamber. The thorough mixing of the evaporated fuel and the airflows prior to combustion results in the production of reduced quantities of the oxide of nitrogen.

Abstract: A turbofan gas turbine engine is provided with a polygonal rigid frame formed by a hub member, a plurality of radially extending spokes and a plurality of straight members connect to the radially outer ends of the spokes. The hub member is secured to the core casing and the straight members are secured to the core casing and the straight members are secured to the fan casing. Brackets are provided at two hard points at the radially outer ends of two adjacent hard points at the radially outer end of two adjacent spokes and brackets are provided at two hard points at the radially inner ends of two other spokes. The spokes have a chord length of six times that of conventional outlet guide vanes. The hard points are connected by links to a bracket at a first axial position on the aircraft stub pylon extension and the hard points are connect by links to a bracket at a second axial position on the stub pylon extension to form a rigid triangular structure to mount the engine from the aircraft.

Abstract: A casing assembly for a ducted fan gas turbine engine comprises a casing around which extend two annular, axially spaced apart, frangible rails. Pieces of cloth woven from an aromatic polyamide fibre are positioned on the casing radially outer surface between the rails. Continuous layers of the cloth are wound around the casing radially outwardly of the rails and cloth pieces. The arrangement provided enhanced containment of any fan blades of the engine which may become detached from the gas turbine engine which carries the casing assembly.

Abstract: A cooling and sealing arrangement in a gas turbine engine comprises a seal assembly (42,30) extending between nozzle guide vanes (20) and combustion chamber discharge nozzles (22). The nozzle guide vanes (20) have platforms (34,36) having extensions (34A,36A) which in one embodiment overlap the downstream ends of the discharge nozzles (22). Cooling air is supplied to the upstream edge portions of the platforms through holes (26,28C) in flanges (24,28) on the discharge nozzles (22). The seal assembly (42,30) defines a chamber (58,70) adjacent the platform extensions which is supplied with cooling air through metering holes (56,68) in the seal assembly. Cooling air exits the chamber (58,70) through holes (34B,36B) in the platform extensions to film cool them.

Abstract: A turbine casing having temperature requirements that vary along the casing and which is at least partially enclosed by a cowling so that a gap is defined between the casing and cowling through which a flow of cooling air may be directed in use, the gap having a magnitude that varies along the casing relative to the temperature requirements of the casing to thereby vary the local velocity of the flow, the cowling being provided with channel shaped portions defining the variations in the gap between the cowling and the turbine casing.

Abstract: A gas turbine engine combustor comprises a combustion chamber the walls of which have a plurality of holes extending therethrough for the flow of cooling air into the interior of the chamber. The holes are so configured and arranged that cooling air is exhausted from them in a flow direction which is normal or oblique to the general direction of gas flow through the chamber and oblique to the portion of the chamber wall local thereto.

Abstract: A gas turbine engine component support assembly (30) comprises a pair of ring members (64) and (74) defining a slot in which a cranked flange (28) extending from the inner surfaces of an array of combustion chamber discharge nozzles (22) is located. One part (28B) of the cranked flange has projections (28C) which engage with radial slots (76) formed in one of the ring members (74) to provide cross-key location of the discharge nozzles. The support assembly (30) is located adjacent the radially inner platform of an array of nozzle guide vanes (20), and also provides support and location for the vanes.