Abstract: A gas turbine engine fan blade (26) comprises a root portion (36) and an aerofoil portion (32). The aerofoil portion (32) has a convex surface (44), a concave surface (42), a leading edge (38) and a trailing edge (40). The leading edge (38) of the aerofoil portion (32) is formed from a harder material than the regions of the concave surface (42) and convex surface (44) immediately adjacent the leading edge (38) such that the leading edge (38) of the aerofoil portion (32) remains pointed. This improves the efficiency of the fan blade (26) and hence the gas turbine engine (10) and reduces flutter of the fan blade (10).

Abstract: A gas turbine engine fan blade (26) comprises a root portion (36) and an aerofoil portion (32). The aerofoil portion (32) has a convex surface (44), a concave surface (42), a leading edge (38) and a trailing edge (40). The leading edge (38) of the aerofoil portion (32) is formed from a harder material than the regions of the concave surface (42) and convex surface (44) immediately adjacent the leading edge (38) such that the leading edge (38) of the aerofoil portion (32) remains pointed. This improves the efficiency of the fan blade (26) and hence the gas turbine engine (10) and reduces flutter of the fan blade (10).

Abstract: A ducted fan gas turbine engine is provided with a fan which is carried by a shaft. A frangible fuse ring maintains the shaft carrying the fan coaxial with the engine longitudinal axis. In the event of severe fan damage, the fuse ring (28) fractures and the fan shaft orbits about the engine longitudinal axis. A pair of threaded members (39, 41), one of which is of generally frusto-conical configuration, cooperate as a result of the orbiting motion to restore the coaxial relationship between the fan shaft and the engine longitudinal axis.

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 variable stator vane arrangement for an axial flow compressor comprises a plurality of stages of variable pitch stator vanes each of which is operated by a respective control ring. The vanes in each stage are connected to the control rings by operating levers. An axially extending member is connected to each control ring by operating links. The axially extending member is arranged in a plane substantially tangential to each control ring and is movable in the plane substantially tangential to the control rings to change the pitch of the variable pitch stator vanes. Hydraulic rams are provided to move the axially extending member. The hydraulic rams may be operated to give either proportional or non-proportional movement of the variable pitch stator vanes. A controller may use signals from position detectors positioned on the control rings or position detectors positioned on the hydraulic rams together with compressor parameters to control the pitch of the vanes.

Abstract: The invention relates to intakes for turbopropeller gas turbine engines of the tractor type. A tractor turbopropeller gas turbine engine comprises a gas generator, an upstream propeller and a downstream propeller. The propellers are positioned upstream of the gas generator, and the propellers are driven via shaft means and gear means. An intake means supplies air to the gas generator, the intake means comprises an intake opening and an intake duct. The intake opening is positioned axially between the propellers, and a portion of the intake duct extends axially through the downstream propeller. This arrangement reduces the length of the shaft and the turbopropeller gas turbine engine, and reduces the weight accordingly. The intake opening is arranged to provide improved protection against ingestion of large foreign objects, such as birds, into the intake duct and gas generator.

Abstract: A control system is provided for the multi-stage axial flow compressor of a gas turbine engine. The control system embraces a stage of variable angle guide vanes, a first detector adapted to detect a first pressure in the compressor which is influenced by the vane angular setting, and a second detector adapted to detect a second pressure independent of the vane setting but bearing a functional relationship to the rotational speed of a compressor of the engine. A control unit is adapted to use the pressures detected by the detectors to cause an actuation mechanism to adjust the angular setting of the guide vanes in a predetermined manner dependent upon the ratio of the second pressure to the first pressure.

Abstract: A rotor blade stage for a gas turbine engine comprises an annular array of aerofoil blades. Each blade is provided with circumferentially extending platforms which are arranged so as to circumferentially spaced apart from the platforms of adjacent blades. Elastomeric dampers are interposed between adjacent blades and are so adapted that upon rotation of the array of blades, each of the elastomeric dampers is centrifugally urged into engagement with the radially inner faces of adjacent platforms of adjacent blades.