Rolls-royce Corporation

Shahrokh Shahpar

Paul Ashley Denman

Christopher L Ford

Emmanuel Aurifeille

Yibin Wang

Greg Carnie

Christopher Armit

Ning Qin

Mark David Taylor

John Coull

Matthew Symes

Joseph R Hardy

Karpaga Vipran Kannan

Andrés E Leiva Gonzalez

Ryan L Rogers

Stefano Caloni

Haidong Li

David Hunt

A fuel swirl nozzle for a gas turbine engine comprises, in axial fuel flow sequence, a fuel inlet portion, a fuel stem portion, and a fuel/air swirler portion. An inlet to the fuel stem portion is fluidly sealed against the fuel inlet portion, and an outlet from the fuel stem portion is fluidly sealed against the fuel/air swirler portion. The fuel/air swirler portion has a cross-sectional profile that is generally dolioform shaped along an axis parallel to the engine's longitudinal axis, and has a first end and a second end. The first end of the fuel/air swirler portion has an air inlet face, and the second end of the fuel/air swirler portion has a fuel/air mixture outlet face. The first end comprises a rim portion enclosing the air inlet face in which the rim portion has a rim portion radius (r) normal to an outer circumference of the inlet face.

Combustor apparatus

A method for shaping an aerofoil by: (a) defining an aerofoil having a nominal shape, the nominal shape defined by; a leading edge, a trailing edge, a root and a tip, a span extending from the root to the tip, a pressure surface and a suction surface extending from the leading edge to the trailing edge; a nominal camber line extending from the leading edge to the trailing edge; (b) defining an edge region on one of the pressure and/or suction surface which extends distance of at least 0.1% but no more than 10% of the camber line length from one of the leading edge or the trailing edge of the aerofoil; and (c) adapting the shape of the pressure and/or suction surface within the edge region such that the edge region of the aerofoil achieves an asymmetric profile with respect to the nominal camber line.

Aerofoil shaping method

Combustion equipment includes annular combustion chamber having an annular upstream wall structure. Annular outer casing surrounds and is spaced from annular combustion chamber and annular inner casing is within and spaced from annular combustion chamber. Ogee shaped annular cowl is positioned upstream of annular upstream wall structure and a stage of compressor outlet guide vanes is positioned upstream of annular cowl. Pre-diffuser extends in downstream direction from stage of compressor outlet guide vanes, pre-diffuser is positioned upstream of cowl. Stage of compressor outlet guide vanes is connected to outer and inner casing. Outer fairing extends radially inwardly from outer casing such that annular outer casing, outer fairing and pre-diffuser define annular outer cavity and outer fairing is spaced from cowl or inner fairing extends radially outwardly from inner casing such that annular inner casing, inner fairing and pre-diffuser define an annular inner cavity and inner fairing is spaced from cowl.

Combustion equipment

A combustion chamber arrangement includes an annular combustion chamber, a plurality of lean burn fuel injectors and a stage of turbine nozzle guide vanes. The stage of turbine nozzle guide vanes is arranged at the downstream end of the annular combustion chamber and includes a plurality of circumferentially spaced vanes extending between and being secured to an annular inner wall and an annular outer wall. The annular inner wall has an outer surface which has a plurality of outer surface portions and each portion is positioned between a pair of circumferentially adjacent vanes and the annular outer wall has an inner surface which has a plurality of inner surface portions and each portion is positioned between a pair of circumferentially adjacent vanes.

Combustion chamber arrangement and a gas turbine engine comprising a combustion chamber arrangement

A computer-implemented method of optimizing an automatically meshable shape is provided. The method includes the steps of: providing a continuous boundary of the automatically meshable shape, the continuous boundary being defined by a spline formed by two or more spline segments, wherein each segment has a terminal point at each end of the segment; parameterizing shape properties of the segments such that selected ones of the shape properties can be varied under operation of an optimization algorithm; linking each segment to its immediately adjacent segments such that the terminal points of the segment remain coincident with the terminal points of its immediately adjacent segments under operation of the optimization algorithm; and using the optimization algorithm to vary selected ones of the shape properties of the segments, whereby the spline defines an adjusted boundary that remains continuous so that the shape remains automatically meshable.

Optimization of an automatically meshable shapes

A method of connecting three-dimensional structured meshes to form a conformal mesh, for solutions of flow problems, including: forming an interface by overlaying surfaces of two structured meshes; defining an interface set of mesh points in the interface; merging mesh points in the interface set that are within a predetermined distance of each other and projecting mesh points of one mesh in the interface that are within a predetermined distance of a line of the other mesh; generating an interface layer in the form of an unstructured mesh; and inserting a node at the center of each hexahedra, and producing an unstructured layer of cells between the two structured meshes by joining the nodes to the mesh points of the interface layer.

Method of connecting meshes

A casing treatment for a fan or compressor stageof a gas turbine engine comprises circumferential groovesA,B,C,D andE which extend in a series disposed between the leading edgesand trailing edgesof bladesof a fan or compressor stage. The groovesvary in depth d in order to optimise the stall margin of the fan or compressor stagewhile preserving peak efficiency.

Turbomachine comprising an annular casing and a bladed rotor

A stator vane assembly for a turbomachine comprising a plurality of circumferentially arranged stator vanes (), the axial position of the stator vanes () and/or the pitch angle circumferentially between adjacent stator vanes () is varied circumferentially around the stator vane assembly. The stator vane assembly reduces the pressure distortion upstream of the fan outlet stator vanes (), reduces the circumferential pressure variation and this reduces blade forced response excitation, noise generation and aerodynamic losses.

Stator vane assembly for a turbomachine

A guide vane is adapted to be mounted in an annular fluid-conveying duct of a gas turbine engine, and has an aerofoil part which extends radially across the duct when the guide vane is thus-mounted. On a transverse cross-section relative to the direction of intended fluid flow across the aerofoil part, at least a portion of the aerofoil part is S-shaped.

Growing community of inventors

Derby, United Kingdom

Shahrokh Shahpar