Rolls-royce Corporation

Gabriel Gonzalez-Gutierrez

Mark J Wilson

Benedict R Phelps

Nigel H S Smith

Marco Barale

Kashmir S Johal

Craig W Bemment

Stephane M M Baralon

Antonios Kalochairetis

Stephane Baralon

Nigel Smith

Domenico Sgro

Benedict Phelps

Jeffrey Green

A gas turbine engineis provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P), is no greater than a certain value. The gas turbine enginemay provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Pressure ratio distributions for a gas turbine engine

A gas turbine engineis provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Gas turbine engine

A fan for a gas turbine engine, the fan comprising a first set of fan blades and a second set of fan blades arranged circumferentially around a hub. Each of the fan blades of the first and second set comprises an organic matrix composite body and a leading edge member connected to the body. The leading edge member of the first set of fan blades has a mass less than the leading edge member of the second set of fan blades.

A fan for a gas turbine engine, the fan comprising a first set of fan blades, each fan blade comprising a root and an axial retention feature provided on the root at a first position, and a second set of fan blades, each fan blade comprising a root and an axial retention feature provided on the root at a second position. Relative to the respective blade on which the respective retention feature is provided, the first position is different to the second position.

A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

Gas turbine engine fan blade

A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

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Derby, United Kingdom

Gabriel Gonzalez-Gutierrez