Abstract: A dual fuel injector for a gas turbine power plant has fuel ducting including gas and liquid fuel ducts which can be removed easily for access to a nozzle of the fuel duct. A separate fluid ducting assembly is attached in the head of each combustor can and has an annular duct, a passage and drillings. The power plant is started on liquid fuel, and runs subsequently on gas produced in a coal gasifier. The flow of air through the duct and passage prevents any unburnt fuel or combustion products from entering the gas duct, and the flow through the drillings prevents carbon build up on face of the injector.

Abstract: For flight safety reasons in gas turbine aeroengines it is desireable that in the event of malfunction of an electronic control system--such as a fuel control system--caused, for example, by overheating due to fire, the control system should protect the engine and the aircraft against the consequences of such a malfunction. In the present invention this is achieved by arranging that the control system is provided with a failsafe unit which monitors the condition of the control system and puts the engine into a failsafe control mode if a control system condition signal generated by the failsafe unit varies beyond a predetermined limit. Thus, to ensure safety under overheating conditions, the control system condition signal is representative of the temperature experienced by the fuel control system and the failsafe control mode is a speed-limiting control mode.

Abstract: The casings of gas turbine engines are subject to loadings which distort the casing and adversely affect the running clearance between rotating and static components contained within the casings. The invention provides an engine having inner and outer housings the inner housing containing one of the engine compressors. The inner casing is mounted on a load transmitting and isolating structure which allows thrust and torque to be transmitted to the outer casing, but isolates the inner casing from distortions of the outer casing. The load transmitting and isolating structure includes a torque reaction and mounting means at the upstream end of the inner casing, a mounting means and a thrust reacting means mounted between the ends of the inner casing. The mounting means are adjacent the bearings of the rotor shaft of the compressor and are not affected to any appreciable extent by outer casing distortion.

Abstract: In a gas turbine engine powerplant which has a requirement for both nozzle outlet area variation and thrust reversal, there is provided a common actuating means in the form of rams which firstly move a cowl portion so as to open a radial gap and thereby effectively increase the nozzle outlet area, and secondly move reverser cascades in a downstream direction into the gap created by moving the cowl portion. A lost motion device is fitted so as to enable achieving of the former movement without affecting the latter movement until such time as the latter movement is also required.

Abstract: It is common practice to pass an ultrasonic probe over a workpiece so as to enable examining of the material for faults such as flaws within the material thickness. Known apparatus for achieving this has a drawback in that the probe supporting structure is such that the probe oscillates about some point within its length, as the structure passes over curved or undulating surfaces. Control over the position of the focal point is thus lost. The invention involves ensuring that the pivot point of the probe coincides with its focal point which in turn enables control of the position of the focal point to be maintained.

Abstract: A turbine provided with at least an annular array of rotary aerofoil blades has a casing, the radially inwardly facing surface of which is provided with a plurality of grooves. The grooves are adjacent the radially outer tips of the aerofoil blades and are so arranged that they direct the gas flow between the blade tips and the casing along the absolute ideal flow path in the region of the blade tips. This ensures that at least some of the momentum of the gases flowing between the blade tips and the casing is transformed into rotating energy as vortices form in the grooves. This energy is subsequently imparted to the tips of the aerofoil blades.

Abstract: A gaseous fuel injector for an industrial gas turbine plant is arranged to operate on fuel produced from a coal gasifier for normal running or natural gas for starting purposes. The injector is self purging to prevent the ingestion of natural gas or combustion products into the passage of the fuel injector for the lower calorific value fuel. The fuel injector has fuel ducts and gas flow passages in a duct assembly attached to the head of each flame tube of a gas generator. For starting, natural gas flows through the duct and the central passage, while air flows through the outer passage preventing ingestion of natural gas and combustion products into the outer fuel duct. When running on fuel from a coal gasifier both ducts run full of fuel, as do the passages. The air for the coal gasifier may be provided by the gas generator.

Abstract: An air intake for a gas turbine engine compressor comprises an arcuate primary duct which is in communication with the compressor of the gas turbine engine and which is provided with frangible wall portion. The frangible wall portion provides a barrier between the primary duct and a secondary duct and is so positioned and adapted as to be impacted and penetrated by a foreign body, such as a bird, which is above a predetermined weight and which has been ingested through the inlet of the primary duct. After penetrating the wall portion the foreign body passes through the secondary duct and is discharged therefrom at a location remote from the primary duct.

Abstract: To minimize secondary losses and flow maldistribution, an aerofoil such as a gas turbine nozzle guide vane, is designed so that in the throat defined by the trailing edge of one aerofoil and the flank of the adjacent aerofoil, the local gas flow streamlines are normal to the trailing edge of the aerofoil at the point of intersection.

Abstract: A fuel system for a gas turbine engine is arranged to schedule fuel flow between primary and secondary fuel injectors in order to obtain good ignition and light round when the engine is operating on diesel fuel. The system includes either two solenoid valves in parallel with flow restrictors or a single solenoid operated valve and two flow restrictors. The two valves or the diverter valve schedule the fuel flow so that at engine start up, the flow to the primary and secondary fuel injectors is according to requirement proportioned. At a certain engine speed the flow to the secondary fuel injectors is reduced to purge flow and the major part of the fuel flow passes to the primary fuel injectors.

Abstract: A control apparatus for the bleed valve of a gas turbine engine compressor enables the engine to be operated with a low risk of compressor surge or stall, particularly on emergency shut down.The control apparatus includes a control valve and a diverter valve. In normal operation the bleed valve is controlled by the control valve which receives a flow of high pressure air from the engine compressor. In an emergency shut down, the diverter valve is operated to allow the high pressure air to flow directly to the bleed valve, without passing through the control valve.The diverter valve can be fuel pressure responsive or solenoid operated.

Abstract: To prevent dust and other debris from entering the compressor (12) of a gas turbine engine (10) an inlet particle separator (30) is provided. Air contaminated with dust and debris is induced into a scavenge chamber (32). An exhaust vent (35) is provided from the scavenge chamber (32) through which the contaminated air is ejected to atmosphere. Flow dividers (40,42,44) are positioned in the scavenge chamber (32) to ensure that the flow into the chamber is uniform around the inlet passage.

Abstract: A containment ring for a gas turbine engine fan duct comprises a rigid cylinder the periphery of which is covered with layers of fibrous material consisting of a plurality of pieces of woven material, around which is wound a further plurality of continuous layers of woven fibrous material which is then covered with a further impervious layer.

Abstract: A method of surface hardening a metal surface in which the surface is firstly coated with a colloidal dispersion of graphite and then locally melted, in an inert atmosphere or vacuum, so that the molten metal reacts with the graphite to form a metal carbide. On resolidification, the metal carbide defines a dendritic carbide structure within the metal adjacent its surface.

Abstract: Carrier particles, e.g., of metal or metal oxide, are doped with a positron-emitting isotope and injected into fluid flowing in a body, such as a gas turbine engine. The positrons decay on coming to rest, giving two gamma rays which are detected, and a 3D image is built up. The particles are capable of stopping a portion of the positions before they escape from the particles into the surroundings; thus the gamma rays indicate more accurately the position at which the positron was emitted, and give a clearer image.

Abstract: A low drag surface construction utilizes a plurality of longitudinally extending, parallel, spaced apart linear vortices extending transversely of the free stream to reduce drag between the free stream and the surface. The surface is provided with stabilizing means which retains the vortices in their relationship with one another but causes them to traverse the surface in the same direction as the free stream but at approximately half the speed. The stabilizing produces a regular variation in boundary flow across the surface and may comprise dynamic means such as sequenced jets of fluid escaping from apertures in the surface.

Abstract: The present invention relates to combustion apparatus for gas turbine engines.An annular combustion chamber of a gas turbine engine comprises an upstream wall formed from a single metal skin. The upstream wall has a plurality of equi-spaced apertures, and a convergent/divergent pot is positioned coaxially with each aperture and extends in an upstream direction therefrom. Each convergent/divergent pot has a radial swirler assembly positioned a its upstream end to supply air into the convergent/divergent pot. A fuel injector is aligned with each convergent/divergent pot and radial swirler assembly to supply fuel into the convergent/divergent pot. A plurality of annular air scoops extend in an upstream direction from the upstream wall, and each annular air scoop is positioned coaxially around a convergent/divergent pot to supply air to the radial swirler assembly. The downstream end of each convergent/divergent pot has apertures for supplying cooling air over the downstream face of the upstream wall.

Abstract: A compound helicopter shown in FIG. 1 of the drawings has wings 12 in addition to a helicopter rotor 14 and has twin powerplants 16 each including a low pressure compressor 18, a core engine 20, a power turbine 22 driven by the core engine and connected through a gearbox 32 to drive the helicopter rotor, and a variable area final, propulsion nozzle 24 which receives the exhaust from the power turbine. Augmentor wing flaps 28 are provided on the wings and fed with air from the low pressure compressor for providing additional lift and thrust from the wings. In operation, at take-off the nozzles 24 are fully opened and all of the power produced by the power turbines 22 is used in driving the helicopter rotor for producing lift. In addition, the augmentor wing flaps 28 are directed downwardly to provide lift. For forward flight the nozzles 24 are closed down, reducing the power to the rotor and slowing it down, while at the same time generating forward thrust.

Abstract: The present invention relates to turbofan gas turbine engines in which the fan assembly is positioned downstream of the core engine of the turbofan gas turbine engine i.e. an aft fan turbofan. Conventional turbofan engines have the fan assembly positioned upstream of the core engine, and the core engine is mounted onto an aircraft pylon. The fan casing is then carried by the core engine. This arrangement in an aft fan would result in losses and intake distortions under side winds and aircraft yaw. The aft fan turbofan has a core engine and a fan assembly which is positioned downstream of the core engine. An upstream and downstream support means, which comprise struts, extend from a fan casing to the core engine and the fan assembly to support the core engine and fan assembly. Upstream mounting means on the fan casing and a downstream mounting means carry the turbofan engine from an aircraft pylon.

Abstract: A propeller module for an aero gas turbine engine comprises two contra-rotating propellers driven by a reduction gearbox positioned axially between the propellers.Pitch change mechanisms operated by a pitch change power unit are provided to change the pitch of the propeller blades. The pitch change power unit is positioned in the hub of the second multi-bladed propeller and is remote from the first multi-bladed propeller. The first and second pitch change mechanisms are mounted on the reduction gearbox. The first pitch change mechanism comprises first pitch change gears which are rotatably mounted on a carrier of the reduction gearbox by spindles which extend axially through planet gears of the gearbox. The second pitch change mechanism comprises second pitch change gears rotatably mounted on the carrier.