Rolls-royce Corporation

Edward S Hadaway

John Frederick Coplin

Brian S Stratford

Norman J Wedlake

John Chew

David R Barker

Howard G Hartland

A pitch change mechanism for a plurality of fan blades (12) comprises adjustment device (60) to rotate the blades (12) about their longitudinal axes. A gear box (34) provides rotary power to drive the adjustment device (60) in response to a control signal. The gear box (34) is mounted in the fan shaft (30) and rotates therewith. An input shaft (33) to the gear box (34) is mechanically connected to a turbine shaft (32) so that the input shaft (33) rotates with the turbine shaft (32). The difference between the speeds of the fan shaft (30) and the turbine shaft (32) is used to drive an output shaft (39) of the gear box (34). The rotary power of the output shaft (34) drives the pitch adjustment device (60) via a plurality of gears (50).

Blade pitch change mechanism

A turbofan gas turbine engine (10) comprises a fan rotor (32) driven by a core engine (11) via a shaft (36) and a reduction gear (38). A lubricant source (63) is arranged to supply lubricant to the reduction gear (38) to lubricate gear teeth and bearings (52) in the reduction gear (38). An annular collector (76) is positioned coaxially around the reduction gear (38) and is secured to the fan rotor (32) and is driven by the reduction gear (38). The annular collector (76) is arranged to collect lubricant leaving the reduction gear (38) on its inner surface. A number of static scoops (82) collect lubricant from adjacent the inner surface of the annular collector (76) to return the lubricant to valves (86). The valves (86) allow the lubricant to return to the lubricant source (63) in normal operation. If the lubricant source (63) fails to supply lubricant to the reduction gear (38) the valve (86) returns the lubricant to the reduction gear (38) to allow the reduction gear (38) to operate for a period of time without a supply of lubricant.

Reduction gear assembly and a gas turbine engine

The present invention relates to air bleed arrangements on a compressor of a gas turbine engine. A rotor of the compressor has a plurality of stages of rotor blades, and the rotor is formed from a plurality of axially adjacent rotor discs. Two axially adjacent rotor discs define a chamber therebetween for a radially inward flow of bleed air. The rotor is provided with a plurality of apertures in rotor disc which bleed air from the compressor and supply the bleed air to the chamber. The adjacent rotor discs have opposing radially extending surfaces, and one of the opposing surfaces has a plurality of circumferentially spaced radially extending vanes and radially extending grooves which direct the bleed air radially inwardly to minimize pressure losses in the bleed air flowing through the chamber. This arrangement is not complex, is relatively easy and inexpensive to manufacture, has reduced weight and allows compressor to be manufactured by welding the rotor discs together.

Compressor and air bleed arrangement

A turbine suitable for a gas turbine engine is provided with a bearing support member which is interconnected with the turbine casing by means of an annular array of aerofoil guide vanes. The whole assembly of support member and aerofoil guide vanes is maintained in a state of tension. The degree of tension in the assembly is controlled so as to be substantially constant throughout the normal operating cycle of the turbine.

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Derby, United Kingdom

Edward S Hadaway