Rolls-Royce Corporation

Stephane M M Baralon

Benedict R Phelps

Mark J Wilson

Craig W Bemment

Christopher Benson

Gabriel Gonzalez-Gutierrez

Nigel H S Smith

Marco Barale

Kashmir S Johal

Neil Wiliam Harvey

Benjamin J Sellers

Michael O Hales

Stuart Andrews

George Crammond

A fan stage of a ducted fan gas turbine engine has a rotor hub having a principal axis of rotation and a plurality of fan blades having a hub end attached to the hub and extending radially towards a tip end so as to define a blade span dimension. Each blade has a leading and a trailing edge, a chord for a section of the blade being a straight line joining the leading and trailing edges within the section. A difference between a stagger angle in a mid-span region and in the vicinity of the tip end of each blade is greater than or equal to 20°. The fan blades are twisted to a greater extent than conventional between the mid-span and tip end. A camber angle difference between the mid-span region and the tip end may be greater than 30 degrees.

Gas turbine engine fan

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.

Gas turbine engine

A gas turbine engine comprises carbon fibre fan blades. At cruise conditions, the fan tip air angle &#x3b8; in the range: 64 degrees&#x2264;&#x3b8;&#x2264;67 degrees. Additionally or alternatively, the fan blade tip angle &#x3b2; is in the range of from 62 to 69 degrees. Arrangements in accordance with the present disclosure provide advantages which may include improved bird-strike performance. This may allow advantages associated with carbon fibre fan blades to be better exploited.

Gas turbine engine having optimized fan

The present disclosure relates to a fan blade for a gas turbine engine, the fan blade comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span. A maximum thickness of the cross-sections through the aerofoil portion from a suction surface of the aerofoil portion to a pressure surface of the aerofoil portion perpendicular to the camber line decreases along the blade span from the root to the tip. There is a discontinuity in the rate of decrease of maximum thickness between a radius at 30% of the blade span from the aerofoil root and 70% of the blade span from the aerofoil root. The rate of decrease of the maximum thickness before the discontinuity is less than the rate of decrease of the maximum thickness after the discontinuity.

Fan blade

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 slot is provided in an endwall of a flow passage, for example between two stator vanes or rotor blades of a gas turbine engine. The length direction of the flow passage is aligned substantially with the main flow through the flow passage. The alignment of the slot means that the 'over-turned' boundary layer flow can be extracted through the slot but with minimal impact on the mainstream flow.

Secondary flow control

A gas turbine engine comprising in flow series: a fan; a series of outlet guide vanes; and a bifurcation. The outlet guide vanes are more densely distributed in a region circumferentially aligned with the bifurcation than in a region circumferentially offset from the bifurcation.

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.

Fan design

A gas turbine engine comprising a fan, a bypass duct positioned downstream of the fan and defined by an inner casing and an outer casing, a series of outlet guide vanes arranged downstream of the fan, the outlet guide vanes extending between the inner casing and the outer casing of the bypass duct; and a bifurcation positioned downstream of the outlet guide vanes and extending between the inner casing and the outer casing. The bifurcation comprises a leading edge, and the leading edge of the bifurcation is shaped so as to protrude axially forward by a varying distance from the inner casing to the outer casing so as to improve uniformity of a static pressure field formed, in use, immediately upstream of the bifurcation.

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Derby, United Kingdom

Stephane M M Baralon