Abstract: A mounting for connecting a turbofan gas turbine engine to an aircraft structure comprises three mounting points at the vertices of a triangle on the aircraft structure and three mounting points on the fan casing of the turbofan gas turbine engine. The first mounting point on the aircraft structure is connected to all of the three mounting points on the fan casing by three links. The second mounting point on the aircraft structure is connected to the second mounting point and the third mounting point on the fan casing by two links. The third mounting point on the aircraft structure is connected to the second mounting point and the third mounting point by two links. The links provide an open spaceframe which defines the structure of a pylon. The mounting has reduced weight compared to conventional mountings and is cheaper to manufacture.

Abstract: A fuel control system for a premixed lean burn combustor in a combustion turbine engine, the combustor having at least two stages (P,S,T), the control system including; an air flow evaluation module having means utilising an iterative algorithm to derive an accurate value of total air flow to the combustor, the evaluation module further having means dividing the derived total air flow value into the proportions (Mp, Ms, Mt) required by the combustor stages (p,s, t) and outputting signals in parallel representing the air flow to respective combustor stages (p,s,t), and a control module associated with each stage (p,s,t) of the combustor and responsible for setting the fuel demand signal for the associated combustor stage (p,s,t), each control module being adapted to set the fuel demand signal with reference to an appropriate one of said output signals.

Abstract: A gas turbine engine is provided with a plurality of compressors and a plurality of turbines. The compressors supply less than half of the oxygen received to the combustor chamber and thus fuel rich combustion occurs in the combustion means. A number of combustion ducts are positioned between adjacent turbines and oxygen is supplied to each of the combustion ducts from the respective compressor via the cooling passages of the respective turbine. Unburned fuel in the exhaust gases of the combustion means is mixed with cooling air in the turbines and is then burnt in the combustion ducts.

Abstract: A root attachment for a rotor blade of a gas turbine engine comprises teeth on the root section of the rotor blade and teeth on the slot of the rotor. The teeth flank surfaces of the teeth on the root section of the rotor blade are arranged at a different angle to the teeth flank surfaces of the teeth on the slot of the rotor. As the load on the rotor blade increases the teeth deflect such that the area of contact between the teeth flank surfaces of the root section and the slot increases to a maximum at the fully loaded condition. The teeth flank surfaces are substantially planar. This enables rotor blade and slot teeth with different stiffnesses to be used for the rotor blades and rotor, for example titanium aluminide turbine blades and nickel base alloy turbine rotor.

Abstract: A multi-stream fuel injector for a gas turbine engine comprises a plurality of concentric member which define a plurality of flow passages carrying HP compressor air or fuel/air mixture. One of the members is formed with a wide-angle, frusto-conical flared lip to propagate a wide-angle fuel/air mixture cone in the combustion region. The lip is of such a cross-sectional configuration that air flowing over the lip surface remains attached to the surface and flows in a non-turbulent manner, initially in a radially inward direction, and subsequently in a radially outward direction. The lip ensures that undesirable regions of recirculation are avoided around the fuel injector.

Abstract: A plurality of hollow blades are produced by assembling a plurality of titanium sheets and a plurality of titanium blocks into a stack. At least one of the sheets is provided with a stop off material on one of its surfaces at a plurality of longitudinally spaced pre-selected areas to prevent diffusion bonding at these areas to define the hollow interiors of the blades. The edges and ends of the sheets are welded together and the blocks are welded to the sheets. The assembly is heated and externally pressurised to diffusion bond the sheets and blocks together. The integral structure so formed is then heated and internally pressurised to superplastically form at least one of the sheets at each of the pre-selected areas to produce a plurality of hollow blades. Thereafter the hollow blades are separated from each other by machining and simultaneously producing the aerofoil shapes.

Abstract: A damper for damping non-synchronous vibration in adjacent shrouded aerofoil blades is in the form of pin which locates in confronting passages in adjacent blade shrouds. The pin is provided with larger diameter portions which are located totally within the passages and frictionally engage the surfaces of the passages to provide vibration damping. The larger diameter pin portions are interconnected by a central, thinner portion. The configuration of the pin reduces the likelihood of it wearing in such a manner that it jams in the passages and no longer provides vibration damping.

Abstract: A dual fuel gas generator (10) is provided with a steam injection system (48) having a bleed-off valve (66) and a steam metering valve which are controlled in concert by a controller (50) to supply steam to a steam manifold (22) for injection into the gas generator at rapidly variable rates independent of the steam supply rate from a boiler (42, FIG. 1). A proportion of the system can be reheated using a steam warm-up flow through the bleed-off valve (66). A further portion of the system can be purged and preheated using compressor delivery air to flow through a steam metering valve (80), a purge valve (74) and a drain valve (94). To control the degree of superheat at the steam manifold (22), water can be injected through a metering valve (90) and a vortex mixer (92), and to purge condensation from a gas fuel manifold (30, FIG. 7 ) of the gas generator, compressor delivery air can be allowed to flow through the manifold (30), a supply line (108) and a purge valve (112 to atmosphere.

Abstract: A preform of a superplastically formed fan blade for a turbofan aeroengine, includes two blade flank components and a core sheet sandwiched between them. A manufacturing process for producing the blade preforms comprises cutting titanium plate material into rectangular prism shaped pieces of suitable length and thickness. Each piece is cut lengthwise along an inclined plane to form two separate tapered panels which become the blade flank components. After machining to a three dimensional profile on the cut faces, pairs of the blade flank components are stacked together with their uncut surfaces confronting each other, the core sheet being placed between them. The blade flank components are then diffusion bonded to the core sheet in a predefined pattern so as to create the blade preform with an internal structure defined by the bonding pattern. The blade preform can then be superplastically formed and finish machined to produce the fan blade.

Abstract: A gas turbine engine with first and second axial flow compressors and an intercooler therebetween is provided with a diffuser to diffuse the air flow leaving the downstream end of the first axial flow compressor at a first radial distance from the axis to the upstream end of the intercooler at a second radial distance from the axis. The diffuser comprises a first radially extending wall, a second radially extending wall and a plurality of diffuser vanes extending therebetween which define a plurality of generally radially extending diffusing passages. The vanes are wedge shaped. This arrangement provides diffusion in a short axial length.

Abstract: A multispool gas turbine engine for an aircraft application includes a transmission system which is operative to transfer power between relatively rotatable engine spools. In a first mode the transmission system is operative to transfer power from the engine's low pressure spool to its high pressure spool and thereby improve the engine's in-flight relight characteristics; in a second mode it is operative to transfer power from the engine's intermediate pressure spool to its high pressure spool to improve engine part speed performance; and in a third mode to transfer power from the engine's high pressure spool to its intermediate pressure spool to assist in ground starting. With each shaft is an associated flow displacement machine (64,66,62) operable as a pump or motor. In other embodiments, a differential is employed for the power transfer between spools.

Abstract: A ducted fan aircraft compound cycle engine comprises a core engine which includes a compressor, three rotary internal combustion engines fed by the compressor and a power turbine to receive the combustion products of the rotary and combustion engines. The rotary internal combustion engines drive the compressor and the power turbine drives the fan. The engine combines the thermal efficiency of the rotary internal combustion engine with the compact size and light weight of the gas turbine engines.

Abstract: A ducted fan gas turbine engine includes a fan which is driven by a low pressure turbine via a load transmitting shaft. The load transmitting shaft additionally drives the accessories of the engine and the aircraft upon which it is mounted. In the event of an engine failure, the power produced by the windmilling of the engine fan is sufficient to provide continuity of drive to the accessories.

Abstract: In operation of an engine it is desirable to be able to keep track of engine condition by measuring the values of various engine parameters such as temperatures and pressures and seeing how the relationship between them change. This must be done while the engine is in a steady state condition, generally this requires that the engine be in a cruise condition for some time but this is impractical for military aircraft. It is suggested that each of the sense parameter values be operated on by a trending algorithm which generates a trend estimate which provides a measure of the rate of change of the parameter value. When all of the parameters have trend values are below some pre-set value the engine condition is taken to be at least temporarily steady state condition and parameter values recorded for later analysis.

Abstract: In a friction bonding operation the friction generating and weld forming forces are usually applied perpendicular to the joint faces. Thus, whenever the joint angle changes to component cassette part of the welding machine must be reset to a new angular orientation. This situation arises in the manufacture of a compressor BLUM (Bladed drUM) for a gas turbine engine, where the blades in successive stages are set at different HADE angles. The invention provides an improved component cassette in which the HADE angle is ignored. The cassette is interlocked with the workplace holder to provide a closed force loop to contain out-of-plane forces while retaining sufficient degrees of freedom to allow the friction generating and welding movements to proceed unimpeded.

Abstract: Friction welding tooling includes a cassette (2,4) for holding a component (6) which is to be welded with great accuracy to a clamped workpiece. The cassette is carried on the distal end of an oscillator arm (8) and urged against the workpiece surface so the cassette has to transmit the frictional heating and weld forming forces to the component. The component (6) must be clamped accurately and securely by the cassette for accurate welding. The cassette comprises a base portion (2) and a floating cap (4). The base portion carried by the oscillator arm (8) has locating features (15,16) to position the component (6) accurately. The floating cap (4) has similar locating features (13,16) but locates entirely on the component (6). The cap (4) is then securely clamped to the base portion (2) with a force transmitting wedge (22) inserted by opposing faces (24,25) in a load path.

Abstract: A ceramic thermal barrier coating layer for a superalloy article is caused to adhere to the superalloy article by applying platinum to the superalloy article and heat treating at a temperature of 1100.degree. C. to 1200.degree. C. for one hour. This causes aluminum to diffuse from the superalloy article into the platinum to form a platinum enriched outer layer which generally includes a platinum enriched gamma phase and a platinum enriched gamma prime phase. An alumina layer is formed between the platinum enriched outer layer and a ceramic coating. The platinum enriched gamma phase and the platinum enriched gamma prime phase in the outer layer reduces the migration of transition metal elements to the ceramic coating to enable a very pure alumina layer to be formed.

Abstract: A combustion chamber head assembly consists of an annular, domed combustor head separated on its downstream side from the combustion region by an annular bulkhead. A number of fuel injector means are spaced apart around the head assembly and extend through to supply fuel-air mixture to the combustor region through apertures in the bulkhead. Fuel is supplied by a nozzle at the upstream end of a mixing region. Air is admitted to this region through a variable geometry airflow arrangement comprising several airflow passage disposed concentrically around the nozzle, the passages may include swirl vanes. A portion of the wall surrounding the mixing region is axially translatable to close-off one of the air inlet passages and direct air from the mixing region into the cavity enclosed by the combustor head and bulkhead from where it escapes into the combustion region through air-only apertures in the bulkhead wall.

Abstract: A vacuum chamber is provided with a pair of seals to enable a fibre to be continuously supplied into and removed from the vacuum chamber for application of a coating. The seals comprise a first chamber which is supplied with argon and a second chamber which is evacuated by a pump. The first chamber is connected to atmosphere by a first hypodermic tube, the first chamber is connected to the second chamber by a second hypodermic tube and the second chamber is connected to the interior of the vacuum chamber by a third hypodermic tube. The hypodermic tubes are coaxial and are telescoped together. The fibre is arranged to pass sequentially through the hypodermic tubes into the vacuum chamber. The argon supplied to the first chamber reduces the air flow down the hypodermic tubes. The seal maintains the pressure in the vacuum chamber and does not contaminate the fibre.

Abstract: The present invention consists in the apparatus for use in welding a first component and a second component together by friction welding, comprising drive means operative to generate reciprocatory movement, a component holder operative to hold said first component, linking means operative to transmit reciprocatory movement from the drive means to the component holder so that in operation said first component performs reciprocatory heat-generating movement, pressure means operative to exert pressure on the component holder so that in operation said first component also performs welding movement, the linking means being such as to accommodate that welding movement during friction welding, and the pressure means being such as to accommodate that heat-generating movement during friction welding.