Rolls-royce Corporation

Benedict R Phelps

Mark J Wilson

Stephane M M Baralon

Gabriel Gonzalez-Gutierrez

Nigel H S Smith

Marco Barale

Kashmir S Johal

Craig W Bemment

Christopher Benson

Benjamin J Sellers

Michael O Hales

Stephane Baralon

Nigel Smith

A gas turbine engine for an aircraft comprises: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a bypass duct delimited by a bypass duct inner wall and a bypass duct outer wall and located radially outwardly from the engine core and downstream of the fan; and an outlet guide vane assembly, located within the bypass duct and, comprising a plurality of outlet guide vanes distributed circumferentially within the bypass duct, each outlet guide vane extending radially along a span between the bypass duct inner wall and the bypass duct outer wall, wherein a space-chord ratio of at least one outlet guide vane, at 50% of the span length from the bypass duct inner wall, is less than 0.72.

Gas turbine engine

A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (&#x3b1;3&#x2212;&#x3b1;1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Low speed fan up camber

A gas turbine engineis provided in which a fan having fan bladesin which the camber distribution relative to covered passage of the fanallows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Gas turbine engine with optimized fan blade geometry

Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.

A gas turbine engine has a cycle operability parameter &#x3b2; in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter &#x3b2; may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Aircraft engine operability

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.

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.

A gas turbine engine has a fan tip air angle and/or a fan blade tip air angle in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of fan tip air angle and/or a fan blade tip air angle may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

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Derby, United Kingdom

Benedict R Phelps