Abstract: An aircraft mounted gas turbine engine system comprises a three shaft propulsive gas turbine engine having three electrical generators; each one respectively associated with one shaft of the engine. Two of the electrical generators are additionally configured so as to function as electric motors so as to facilitate power transfer between the engine shafts. One of the electric generators provides the primary source of electrical power for the aircraft, while the remaining electrical generators provide electrical power for the engine and back-up power for the aircraft.

Abstract: A double annular combustor for a gas turbine engine is provided with annular arrays of main and pilot fuel injection modules. The main fuel injection modules are of the premix type so as to vaporise fuel. However, the pilot fuel injection modules are configured so as to function as both premix and airspray fuel injectors. During starting and low power conditions, the pilot fuel injectors are operational alone in their airspray mode. However during high power conditions, both the main and pilot fuel injection modules function as premix injectors. The arrangement reduces noxious emissions.

Abstract: A bladed rotor has blades fitted in grooves around its rim. The blades are prevented from making significant movement lengthways of their grooves by abutment members which are so formed as to enable fitting and removal without the use of special tooling.

Abstract: In a ducted fan gas turbine engine the fan module is connected rigidly to aircraft structure and flexibly to the core gas generator via flanged annular diaphragms (62,66) thus obviating the exertion of bending loads on the core gas generator casing when the fan intake experiences asymmetric aerodynamic loads.

Abstract: A ducted fan gas turbine engine is mountable on an aircraft structure via a pylon which includes a yoke. The yoke straddles the compressor module of the engine and is connected thereto at points diametrically opposite each other, thus obviating thrust links and the need for radial struts in the plane of the yoke.

Abstract: A rotor disc (15) has an axially enlarged hub portion (18) with a bore (20) through which a turbine shaft (19) extends in a coaxial, radially spaced apart relationship, the annular gap between the hub portion (18) and the shaft (19) receives a flow of temperature regulating fluid. The shaft (19) is provided with axially extending elongate projections (22) to increase the turbulence and relative velocity of the temperature regulating fluid adjacent the projections (22).

Abstract: Variable amplitude, linear reciprocation of an output ram is produced by converting rotary motion into linear motion using a swash plate connected by a bending element to drive a crank. The rotary prime mover and swash plate are mounted on a pivotable frame so as the axial throw of the swash plate is determined by the angular offset of the rotational axes of the swash plate and crank. When the axes are coincident the throw, and therefore the ram stroke, is zero but progressive angular displacement produces a progressively increased stroke. Thus, the frequency of reciprocation is determined by the angular speed of the prime mover, and the stroke is controllably variable between maximum and zero.

Abstract: The frictional loads involved in a friction welding operation, for joining metal parts, may be several tons and the mass of the welding apparatus is correspondingly large. Thus, in addition to the energy which is converted into heat at the weld interface a substantial amount of work has to be done simply to overcome the inertial loads of the apparatus. An inertial load cancelling system is proposed in which a double acting gas spring arrangement is coupled to the welding oscillator output so that the work done in overcoming the inertia of the apparatus is returned to the system when the oscillator motion is reversed at the ends of each stroke. Thus, after the initial system input the oscillator has only to supply work to overcome internal frictional and that converted to frictional heat at the weld joint interface.

Abstract: A superalloy turbine blade (10) has a MCRALY bond coating (22) and a thermal barrier coating (24). The thermal barrier coating (24) comprises a plurality of alternate layers (25, 27) which have different structure to produce a plurality of interfaces (26) substantially parallel to the metallic substrate (20)/ bond coating (22) interface. These interfaces (26) provide paths of increased resistance to heat transfer to reduce thermal conductivity of thermal barrier coating (24). The structure in the alternate layers (25, 27) of the thermal barrier coating (24) is columnar (28), or segmented, to ensure that the strain tolerance of the thermal barrier coating (24) is not impaired. The columnar grains (28) in the layers (25, 27) extend substantially perpendicular to the metallic substrate (20)/ bond coating (22) interface. This enables thermal barrier coating (24) of reduced thickness and weight to be used.

Abstract: A gas turbine engine comprises turbine and compressor elements on a rotary shaft. The shaft is rotably mounted on preloaded axial thrust bearings. As it rotates in use the turbine and compressor elements create an axial thrust load on the bearings. According to the invention that axial thrust load is opposed by an electromagnetic coil which is energised to exert an axial force on the shaft in opposition to the axial thrust load and preferably completely balancing the axial thrust load. Bearing life is much extended.

Abstract: A plurality of heat resistant tiles are provided which combine to form an internal liner around the walls of an annular combustor. In a first embodiment a plurality of tiles are arranged in an overlapping relationship to form a fully annular shield around an upstream bulkhead wall. In a second embodiment a plurality of tiles are arranged row by row in an overlapping relationship to form a protective shield around the combustors radially spaced sidewalls. In both embodiments the tiles are attached to the combustor in such a manner that at least one edge of the tiles is clamped against the combustor by an overlapping portion of a neighboring tile. The underside of the tile is exposed to a high pressure flow of cooling air which issues as a film over the exposed surface of the tile. The underside of each tile is sealed by virtue of the clamping effect holding the overlapping edges of adjacent tiles in sealing engagement.

Abstract: A brush seal element (14) comprises a mounting ring (20) that carries radially inwardly directed bristles (21). A backing ring (23) attached to the mounting ring (20) is positioned alongside the bristles (21). A chamber (25) defined by the bristles (21) and the backing ring (23) is pressurised with fluid to counteract the lateral forces imposed upon the bristles (21) by a region of high fluid pressure that they are operationally adjacent. As a result of this balancing of forces the bristles (21) are free to move radially inward and outward so that their free ends (22) remain in sealing engagement with an adjacent surface.

Abstract: An annular heat exchanger suitable for a gas turbine engine comprises a first continuous sheet of material arranged in a spiral and a second continuous sheet of material arranged in a spiral. A first axially extending passage is defined between the first surface of the first sheet and the third surface of the second sheet and a second axially extending passage is defined between the second surface of the first sheet and the fourth surface of the second sheet. The first passage is closed at its axial ends by spiral seals between the first sheet and the second sheet. The second passage is open at its axial ends. A plurality of radially extending passages are provided to supply a first fluid at a first end into the first passage, to remove the first fluid from a second end of the first passage and to interconnect adjacent turns of the first passage. The passages extend radially through the second passage.

Abstract: A gas turbine combustion chamber which has primary, secondary and tertiary combustion zones in flow series has a secondary mixing duct and a tertiary mixing duct. The secondary and tertiary mixing ducts reduce in cross-sectional area from their intakes to their outlet apertures to provide an accelerating flow through the mixing ducts to prevent the formation of recirculating zones. Fuel injectors have fuel discharge apertures downstream of any recirculating zones formed at the intakes. The fuel injectors extend across a major portion of the width of the ducts to effectively subdivide the ducts over at least part of the streamwise length of the ducts. The portions of the fuel injectors within the ducts are oval shaped in cross-section and the portions outside the ducts are aerofoil shaped in cross-section. The fuel injectors reduce in dimension perpendicular to the widthwise direction of the duct.

Abstract: An abradable seal comprises a plurality of hollow aluminosilicate spheres in a matrix of aluminium phosphate. The matrix of aluminium phosphate also has an aluminosilicate filler. The hollow aluminosilicate spheres have a diameter in the range of 400 to 1800 microns, preferably a diameter in the range of 800 to 1400 microns. The weight proportion of hollow aluminosilicate spheres in the abradable seal is 30% to 50%. The density of the abradable seal is 1.5 grammes per cubic centimetre. The hollow aluminosilicate spheres in the aluminium phosphate matrix have a relatively high temperature capability of approximately 1300.degree. C. maximum. This makes the abradable seal better able to withstand the environment in the turbine of a turbofan gas turbine engine.

Abstract: A gas turbine engine turbomachine (10) comprises a turbine section (18) in which cooling air is supplied from a compressor section (14) via a pipe (68) to a chamber (60). The chamber (60) supplies the cooling air through passages (58) in the nozzle guide vanes (30) to impinge upon a stator disc (32). A valve (70) in the pipe (68) is used to switch the flow of cooling air on and off. The valve (70) is opened to allow the cooling air to cool the stator disc (32) to shrink the stator disc (32) to minimise clearance between a shroud (34) supported by the nozzle guide vanes (30) and the turbine blades (28) during one mode of operation. The valve (70)is closed to allow the stator disc (32) to expand to increase the clearance between the shroud (34) and the turbine blades (28).

Abstract: A laser barrier material for preventing far-wall strike during a laser drilling operation comprising a finely divided crystalline material, preferably a metal salt which is thermally stable and possesses a high melting point such as sodium chloride. The barrier material may be introduced into the interior of a component by pouring through an aperture or by making it into a paste with water and injecting it. The material is then subsequently removed by washing the component with water.

Abstract: A cast and hot isostatically pressed gamma titanium aluminide turbine blade (26) has an aerofoil portion (38) including a concave pressure surface (44) and a convex suction surface (46). The concave pressure surface (44) has surface irregularities (48), produced by the action of the hot isostatic pressing process on voids (56) located within the cast turbine blade (26), located in that region (50) of the concave pressure surface (44) where in operation aerodynamic separation (G) occurs to minimise the aerodynamic effects of the surface irregularities (48) upon the operation of the turbine blade (26). The remaining portion of the concave pressure surface (44) and the whole of the convex suction surface (46) are substantially free of surface irregularities (48).

Abstract: A coated article includes a superalloy substrate, an intermediate bond coat and a thermal barrier coating. The bond coat may include a platinum aluminide layer underlying a thin oxide layer. The thin oxide layer may include alumina. The coated article has high strength and durability at high temperatures over extended periods of time and thus is especially useful in the form of, e.g., a turbine blade or turbine vane.

Abstract: In a gas turbine engine including a turbine which has a closed cycle steam cooling system and an open cycle gas cooling system sealing arrangements are provided between adjacent pairs of turbine rotor stages. The sealing arrangement has axially extending passages to supply cooling gas from one of the turbine rotor stages to the cooling gas passages in the turbine rotor blades. Each sealing arrangement comprises a pair of sealing formations each made up of a plurality of sealing segments. Also the sealing arrangement has axially extending passages to supply steam from one of the turbine rotor stages to the steam cooling passages in the turbine rotor blades.