Abstract: A turbine blade tip clearance control system has a rigid two part outer casing (42) which sandwiches a control ring (48) therebetween, and an air pressurised flexible inner casing (28) which carries shroud segments (22) within it. Struts (40) span the annular space between the casings (42, 28) and prevent flexing of casing (28) until blade tip clearance needs adjusting, whereupon, ring (48) is heated, along with the adjacent portion of outer casing (42) and expands, allowing casing (28) to flex outwards, thus lifting the shroud segments (22) away from the blade tips (24). Closure of the tip clearance is achieved by cooling ring (48), the resulting contraction thereof, via the struts (40), flexing the inner casing (28) and shroud segments (22) inwards, against the air pressure.

Abstract: A turbine blade tip clearance control system has a rigid two part outer casing (42) which sandwiches a control ring (48) therebetween, and an air pressurised flexible inner casing (28) which carries shroud segments (22) within it. Struts (40) span the annular space between the casings (42, 28) and prevent flexing of casing (28) until blade tip clearance needs adjusting, whereupon, ring (48) is heated, along with the adjacent portion of outer casing (42) and expands, allowing casing (28) to flex outwards, thus lifting the shroud segments (22) away from the blade tips (24).

Abstract: Use of an optical fiber for the direct receipt of heat radiation for transmission to a remote pyrometer is enabled by the provision of an apertured, contaminant free compartment in the component being heated, and aligning the heat receiving end of the optical fiber with the aperture so as to receive radiated heat from within the compartment.

Abstract: Use of an optical fibre for the direct receipt of heat radiation for transmission to a remote pyrometer is enabled by the provision of an apertured, contaminant free compartment in the component being heated, and aligning the heat receiving end of the optical fibre with the aperture so as to receive radiated heat from within the compartment.

Abstract: A ducted fan gas turbine engine is provided with a fan which is carried by a shaft. Frangible fuse pins maintain the shaft carrying the fan coaxial with the engine longitudinal axis. In the event of severe fan damage, the fuse pins fracture and the fan shaft orbits about the engine longitudinal axis. A spoked resilient member is provided to exert a radially inward restoration force upon the fan shaft following such an event in order to reduce vibration.

Abstract: In a gas turbine engine, compressor delivery air is fed into air inlets in hollow turbine nozzle guide vanes, and then ejected into the gas stream via holes in the vane walls. Valving is provided so as to provide modulation of the cooling airflow over and above the modulation achieved by compressor output when it changes with engine power setting. Excessive cooling of the gas stream is thus prevented.

Abstract: When cooling air from a gas turbine engine compressor is supplied in greater quantities than is needed, the flow thereof is reduced firstly and mainly by ensuring that the air leaves a blade stage cooling thereby at a radius less than that of its entry. A pressure head results which slows the cooling air flow rate prior to its entering a plenum chamber defined by the disc of the stage and fixed structure. There is a reduced tendency of the air to enter the gas stream.

Abstract: In an annular array of stator aerofoil vanes (25 ) each of the vanes (25 ) has a flank portion (35 ) which is pivotally attached to the remainder of its associated vane (25). Actuation means (60) are provided to pivot each flank portion (35) relative to the remainder of its associated vane (25 ). Such movement of the flank portions (35 ) provides variation in the cross-sectional areas of the throat (61) defined between adjacent stator vanes (25).

Abstract: An aerofoil vane actuation mechanism suitable for a gas turbine engine (10) comprises an annular array of pivotable aerofoil vanes (21) which are linked to an annular actuation member (30) by a plurality of levers (27). Each aerofoil vane (21) additionally has an indication lever (31) attached to it. Each indication lever (31) cooperates with the actuation member (30) so that under normal conditions, during pivoting of the vanes (21), no load transfer takes place between them. However in the event of one or more of the vanes (21) failing to pivot properly, such load transfer does take place. Means (37) are provided to detect such load transfer to provide an indication of such failure of a vane (21) to pivot properly.

Abstract: A cooled aerofoil blade for a gas turbine engine has a hollow trailing edge region in which bridging members interconnect the convex and concave flanks of the blade. The bridging members are each provided with angled cooling passages which are in operation exposed to a cooling air flow and direct that flow onto the internal surfaces of the convex and concave flanks so as to provide impingement cooling thereof.

Abstract: A variable guide vane arrangement includes a control ring located in a fixed axial position and an operating lever having a first end and a second end which is secured to the control ring with the first end positioned between two flanges and secured thereto by a pin; the second end is connected to a portion which fits into a bore and is slidable with respect thereto to vary the length of the operating lever.

Abstract: The invention relates to variable guide vane arrangements for axial flow compressors of aircraft gas turbine engines. The variable guide vane arrangement comprises a plurality of stator vanes rotatably mounted in a stator structure of the compressor. A control ring surrounds and is normally coaxially with the compressor axis, and a plurality of operating levers extend from the control ring to their respective stator vane. The control ring is movable laterally with respect to the axis of the compressor so that the stator vanes in a first half of the compressor are rotated in one direction so that the first half of the compressor operates at a higher pressure ratio and the stator vanes in a second half of the compressor are rotated in the opposite direction so that the second half of the compressor operates at a lower pressure ratio.

Abstract: In a multi stage ducted fan or ducted propeller gas turbine engine, a fan flow exit path is provided in the cowl between axially adjacent stages of blades. The stage of blades downstream of the exit are each pivotable to attitudes which on continued rotation of the stage, induces a reverse flow of ambient air into the duct via the nozzle. The normal and reversed airflows meet under the exit path and pass therethrough to atmosphere. The need for blocker doors in the duct is obviated and energy is imported to the fan air passing through the exit by both stages of blades.

Abstract: A gas turbine engine is disclosed in which a particularly efficient use of cooling air is made. In this engine, at least part of the cooling of the turbine blades is effected by a bleed of air from one of the compressors. This bleed air is ducted to flow between the shanks of adjacent rotor blades to cool them, and the used bleed air then rejoins the main gas flow of the engine upstream of the turbine rotor.

Abstract: A rotor blade for a gas turbine engine comprises an aerofoil, shank portion and a root by which the remainder of the blade may be supported from a rotor disc. Cooling for the aerofoil is provided by passages in the leading portion of the aerofoil which are fed with cooling air from an entry aperture and a sealed liquid cooling circuit which contains a cooling liquid. This circuit comprises a liquid feed passage adjacent the trailing edge of the aerofoil and connected adjacent the tip of the aerofoil with a vapor return passage, both passages communicating with a sealed cavity within the shank of the blade. In operation the cooling fluid (e.g. metallic sodium) flows under the influence of centrifugal and other forces along the feed passage, is vaporized and returns to the cavity via the return passage. The cavity acts as a condenser for the fluid.

Abstract: A turbine for a gas turbine engine is provided with a rotary stage comprising an annular array of rotor blades surrounded by an annular shroud member. The shroud member is hollow, the hollow interior being provided with a heat transfer medium which, in operation, functions as a condensable vapor and liquid whereby the shroud member is a heat pipe. The shroud member is further provided with a wall adjacent the aerofoil blade which is adapted to be penetrable by a detached blade or blade portion. The wall of the shroud member is dimensioned so that any such detached blade or blade portion which penetrates the wall enters the hollow shroud ring and is retained within it.