An airfoil assembly for a turbine engine comprising an outer band, an inner band radially spaced inwardly from the outer band to define an annular region, and multiple airfoils circumferentially spaced within the annular region. Each corresponding airfoil of the multiple airfoils can project from a surface at a root and can further include an outer wall defining a pressure side and a suction side. A projection can extend upwardly from the surface on the pressure side and a valley can extend into the surface on the suction side to define a contour in the surface.

A vane for use in a gas turbine engine includes an airfoil having a suction side and a pressure side, and a transpirational flow acoustic liner disposed in the airfoil. The liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet, and a backing sheet coupled to the segmented member the segmented member such that the segmented member is positioned between the face sheet and the backing sheet. The segmented member includes a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures. The backing sheet has a plurality of second apertures.

An aerofoil assembly includes a platform and one or more aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. The one or more aerofoils are disposed between the first edge and the second edge. Each of the one or more aerofoils has a leading edge proximal to the first edge and a trailing edge distal to the first edge. The platform defines one or more recesses disposed between the leading edge of each of the one or more aerofoils and the first edge.

A steam turbine has a diffuser that is configured to guide steam to an outside of a casing. The diffuser has an outer guide that gradually expands to an outer side in a radial direction and an inner guide that is disposed at intervals to an inner side in the radial direction with respect to the outer guide. The inner guide has an inner curved diameter-expanded portion that gradually expands to the outer side in the radial direction while being curved from the first side to the second side in the axial direction. The outer guide has a first diameter-expanded portion that gradually expands to the outer side in the radial direction with a first radius of curvature, and a second diameter-expanded portion that gradually expands to the outer side in the radial direction with a second radius of curvature larger than the first radius of curvature.

Methods and apparatus for reducing flow distortion at engine fans of nacelles are disclosed. An example apparatus for reducing flow distortion at an engine fan of a nacelle includes a plurality of nozzles radially spaced about an inner wall of the nacelle. In some examples, respective ones of the nozzles are positioned to eject corresponding respective jets of fluid adjacent the inner wall in a downstream direction toward the engine fan. The example apparatus further includes a controller to selectively activate the respective ones of the nozzles according to a time-based sequence. In some examples, the time-based sequence corresponds to a directional sequence that moves in an arcuate direction along a circumference of the inner wall.

A turbine engine assembly including an air flow guide assembly, including at least one guide vane and at least one structural arm, the vane and arm extending radially about an axis. The arm includes an upstream end portion having a guide vane profile and including a leading edge aligned with that of the vane; a downstream portion; and an intermediate portion including an upper surface extending between an upstream end point and a downstream end point. The upstream end point is separated from the leading edge of the arm by an axial distance of between 0.2c and 0.5c, c being the length of the axial chord of the vane, and the angle of the tangent to the upper surface at the upstream end point is equal to that at the downstream end point ±1 degree.

A blade of a fan or compressor that reduces loss by enlarging a laminar flow region over a blade surface is provided. The blade is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. A blade surface angle change rate is based on an angle formed by a tangent to the blade surface and the axis of the engine, the leading edge blade surface angle, and the trailing edge blade surface angle at. In each of the subsonic region and the transonic region, values of the blade surface angle change rate on the pressure and suction surfaces are defined at predetermined axial locations along the chord on the pressure and suction surfaces.

A vane for a turbine engine turbine includes a blade and at least one platform that radially extends the blade, and a system for reducing vortices including at least one circuit including a duct extending from at least one air intake orifice formed in the platform upstream of the leading edge as far as at least one air exhaust slot formed on the trailing edge.

An aerofoil assembly includes a platform and one or more aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. The one or more aerofoils are disposed between the first edge and the second edge. Each of the one or more aerofoils has a leading edge proximal to the first edge and a trailing edge distal to the first edge. The platform defines one or more recesses disposed between the leading edge of each of the one or more aerofoils and the first edge.

An aerofoil assembly includes a platform and a plurality of aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The platform defines one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.