Abstract: A composite nozzle segment includes one or more airfoils extending radially between arcuate radially outer and inner band segments having band plies curved about a centerline axis, an outer chordal seal line on an outer aft flange of the outer band segment, and thickened areas in corners of the outer aft flange along pressure and suction side edges of the outer band segment. The thickened areas may be thickest at the pressure and suction side edges and thinnest near a middle of the outer band segment. The thickened areas taper off from an aft end of the thickened area towards a forward flange of the outer band segment. Insert plies may be disposed between some of the band plies within the thickened areas may be grouped together in a discrete group or be interspersed with some of the band plies. The insert plies may have triangular shapes.

Abstract: A turbocharger having a variable nozzle with stepped two-stage vanes (50), the variable nozzle comprising a tubular piston (70) disposed in the bore (44) of the turbine housing (38) such that the piston (70) is axially slidable adjacent to the vanes (50) that extend across the nozzle. Each vane defines a first vane stage (50a) proximate the free end of the vane, the second vane stage (50b) having a different aerodynamic contour in comparison with the first vane stage (50a), each vane comprising a step (60) transitioning from the first vane stage (50a) to the second vane stage (50b). The piston (70) in a closed position closes the second vane stage (50b) so that exhaust gas flows only through the first vane stage (50a). The second vane stage (50b) is progressively opened as the piston (70) is axially slid toward an open position.

Abstract: A vane cluster has a coated metallic substrate. The cluster includes a platform and a shroud. At least first and second airfoils extend between an outer face of the platform and an inner face of the shroud. Each airfoil has a pressure side and a suction side. The pressure side of the first airfoil faces the suction side of the second airfoil. The cluster includes a cooling passageway system including one or more first feed passageways in the first airfoil and one or more second feed passageways in the second airfoil. At least a first side selected from the pressure side of the first airfoil and the suction side of the second airfoil includes a first region with a local thinning or gap in the coating. Along the first side, the cooling passageway system includes means for locally cooling said first region.

Abstract: A steam turbine singlet nozzle airfoil with integral outer sidewall is engaged with an inner ring and an outer ring in a nozzle assembly. The interface of the outer sidewall with the outer ring may include a plurality of mechanical hooks on one or both of the upstream face and the outer radial face of the outer sidewall that engage with complimentary structures on the outer ring. The outer interface may further include low energy welds along limited distances of one or both of the upstream or downstream interface of the outer sidewall and the outer ring. An inner radial end of the singlet nozzle airfoil is pinned into position and fastened to the inner ring. Without a need for high heat welds, distortion of the airfoil and the steam flow path and the associated rework is eliminated and stage performance is improved.

Abstract: A circumferential row of aerofoil members span an annular duct for carrying a flow of compressible fluid. The duct is centered on the axis of a turbomachine. Pressure and suction surfaces of neighboring aerofoil members bound respective sectoral passages which receive the flow of compressible fluid. The row of aerofoil members includes at least one radial endwall to which the tangent to the trailing edge of each aerofoil member at midspan is substantially orthogonal. Each aerofoil member has reverse compound lean. Each aerofoil member further has a leading edge which has a position at the endwall which is upstream of its position at midspan. The endwall at each sectoral passage has a non-axisymmetrical cross-section formed by a region that, in meridional section, is convexly profiled immediately adjacent the pressure surface and a region that, in meridional section, is concavely profiled immediately adjacent the suction surface.

Abstract: An axial-flow gas turbine engine includes a plurality of inlet guide vanes (V) which are radially disposed in an annular gas passage defined between an inner peripheral wall (Ch) and an outer peripheral wall (Ct) of a turbine. The inner peripheral wall (Ch) of the gas passage includes inner peripheral concave portions (Cc1 and Cc3) on an upstream side, and inner peripheral convex portions (Cv1 and Cv3) on a downstream side. The outer peripheral wall (Ct) of the gas passage includes outer peripheral convex portions (Cv2 and Cv4) on an upstream side, and outer peripheral concave portions (Cc2 and Cc4) on a downstream side. Therefore, a pressure difference in a radial direction of the inlet guide vane V is reduced or partially reversed, and a secondary flow toward an inner side in the radial direction can be suppressed to reduce pressure loss.

Abstract: In one aspect, a method for assembling an outlet guide vane is provided. The method includes providing an outlet guide vane frame having a first side and a second side. The outlet guide vane includes a radially inward flange coupled to a radially outward flange by a leading edge. A trailing edge is coupled to the radially inward flange and the radially outward flange aft of the leading edge. A cavity is defined between the radially inward flange, the radially outward flange, the leading edge, and the trailing edge. A first and a second mating surface circumscribe the cavity on the first and second sides respectively. The method further includes coupling a filler portion within the cavity wherein the filler portion includes a third side and a fourth side, coupling a first skin to the third side and the first mating surface, and coupling a second skin to the fourth side and the second mating surface, wherein at least one of the first skin and the second skin are fabricated from a composite material.

Abstract: A first insert (15) forming a ring is placed inside and in contact with a core (10) in degradable material. On the outer face of the core a plurality of inserts (20) is arranged each comprising a platform (21) applied against the outer face of the core and at least one blade (22) extending substantially radially outwards. A fibrous structure (56) is wound around the core passing over the platforms (21) and over the inner surface of the ring-shaped insert (15). Molding of the fibrous structure, impregnated with a composition containing a resin, is performed before polymerization of the resin and removal of the core.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE A. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches. The profile sections at the Z distances can be joined smoothly with one another to form a complete airfoil shape.

Abstract: A vane for a turbine assembly of a turbocharger includes a first airfoil that includes a length between a leading edge and a trailing edge, a second airfoil that includes a length between a leading edge and a trailing edge where the length of the first airfoil optionally differs from the length of the second airfoil, and one or more intra-vane throats defined at least in part by the first airfoil and the second airfoil. Various other examples of devices, assemblies, systems, methods, etc., are also disclosed.

Abstract: A gas turbine engine assembly has combustion gases flowing through a gas flow path. The gas turbine engine assembly includes a stator assembly comprising a stator vane that extends into the gas flow path; and a turbine rotor assembly downstream of the stator assembly and comprising a turbine platform and a turbine rotor blade extending from the turbine platform into the mainstream combustion gases flow path. The turbine rotor blade includes a pressure side and a suction side opposing the pressure side that extend from a leading edge to a trailing edge. The combustion gases form horseshoe vortices at a formation area adjacent the leading edge of the turbine rotor blade, and the turbine rotor assembly further includes a first set of holes in the turbine platform for directing first jets into the formation area of the horseshoe vortices.

Abstract: A two-stage high pressure turbine includes a second stage blade having an airfoil with a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Abstract: A blade for a stator in a torque converter including a first face, which includes a first step, and a second face located substantially circumferentially opposite the first face. In one embodiment, the first face includes first and second disjointed or misaligned segments, and the first step connects the first and second segments. In another embodiment, the second face includes a second step. In yet another embodiment, the second face includes third and fourth disjointed or misaligned segments, and the second step connects the first and second segments.

Abstract: An axial flow fan includes an impeller rotatable about a central axis with a plurality of rotor vanes, a motor that drives the impeller, a base portion that supports the motor, a housing that includes an intake vent, an exhaust vent, and an inner peripheral surface surrounding the impeller and the motor, and a plurality of stator vanes that respectively connects the base portion and the housing, wherein the inner peripheral surface includes a first inner peripheral surface arranged to increase a distance from the central axis toward the intake vent or the exhaust vent in an axial direction, and a recess provided between the first inner peripheral surface and a stator vane included in the plurality of stator vanes and facing the first inner peripheral surface. Thus, airflow is allowed to smoothly pass through the housing, resulting in a decrease in noise generated in the fan.

Abstract: An axial flow fan according to the present invention comprises an impeller which rotates about a central axis and including a plurality of blades, a hollow member accommodating therein the impeller, a base portion which is arranged at the hollow member and supports the base portion in a rotatable manner, a plurality of inner air guide members each connected to the base portion, and a plurality of outer air guide members each connected to the hollow member. The inner air guide members and the outer air guide members each include a first edge member and a second edge portion. A length of the first edge member and that of the second edge member of the outer air guide member are greater than those of the first edge member and the second edge member of the inner air guide member.

Abstract: The present application provides an axial flow turbine. The axial flow turbine may include a stator casing and a turbine bucket positioned about the stator casing. A tip shroud may be positioned on the turbine bucket. A shroud tail may be attached to the tip shroud at a downstream end of the tip shroud.

Abstract: In a transonic region with a Reynolds number not more than a critical Reynolds number, a flow velocity distribution on an extrados of an airfoil has a single supersonic maximum value within a range of up to 6% from a leading edge on a chord, or a shape factor has a maximum value in a region of 6 to 15% from the leading edge on the chord, the value being nearly constant in a region of 30 to 60% and gradually can increase up to 2.5 in a region downstream of 60% of chord. A pressure loss in a low Reynolds number region can be drastically reduced, while conventionally keeping low the pressure loss in a high Reynolds number region. Moreover, this pressure-loss reduction effect in the low Reynolds number region is exerted even if an inflow angle is changed in a wide range.

Abstract: A sectorized nozzle for a turbomachine, the nozzle being made up of cylindrical sectors and comprising two annular platforms interconnected by substantially radial vanes, together with an annular rail for supporting elements of abradable material, each inner platform having longitudinally-extending edges that are V-shaped, and each rail sector having, on at least one of its sides, means for bearing axially on corresponding means provided on an adjacent nozzle sector.

Abstract: A stator vane for a gas turbine includes a vane airfoil which extends in the longitudinal direction of the vane and which is delimited by a leading edge and a trailing edge, and also an outer platform, the inner side of which is exposed to the hot gas which flows through the gas turbine, and on which provision is made for a hook-like fastening element, projecting outwards in the region of the trailing edge, for fastening the stator vane on a casing of the gas turbine, which fastening element, on its side facing the trailing edge, has a locating slot above the trailing edge for the fixing of a heat shield which adjoins the outer platform of the stator vane in the flow direction of the hot gas. Provision is made on the outer platform of the stator vane between the locating slot and the trailing edge of a structure for reducing the thermal and mechanical stresses in the region of the transition between trailing edge and outer platform.

Abstract: Systems and methods for integrating heat exchanger elements of HRSG systems with gas turbine exhaust diffusers are provided in the disclosed embodiments. The systems and methods may include integrating heat exchanger elements, such as steam pipes, with various components of an exhaust diffuser. For example, the heat exchanger elements may be integrated with inlet turning vanes, exhaust frame struts, exit guide vanes, associated support structures, and other components of the exhaust diffuser. In addition, the heat exchanger elements may be integrated with multiple components of a single exhaust diffuser. Moreover, the heat exchanger elements may be integrated with the components of the exhaust diffuser within an airfoil, which may encompass both the heat exchanger elements and the individual component of the exhaust diffuser. The use of airfoils may help ensure certain aerodynamic properties of the heated exhaust gas flowing across the exhaust diffuser components.

Abstract: A bucket for a turbine is described. The bucket includes a dovetail portion configured to couple the bucket to a turbine wheel, the dovetail portion having a lower surface. The bucket also includes a shank portion that extends from the dovetail portion and an airfoil having a root and a tip portion, an airfoil shape, and a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table I. The bucket also includes a plurality of cooling passages. The plurality includes no more than five cooling passages that extend between the root and the tip portion of the airfoil. Each of the cooling passages exits the airfoil at the tip portion, the plurality of cooling passages are positioned in a camber line pattern.

Abstract: A gas turbine vane to improve vane performance by addressing known failure mechanisms. A cooling circuit to the trailing edge of a vane airfoil is fed from the outer diameter platform, which prevents failure due to an oxidized and eroded airfoil trailing edge. The gas turbine includes an outer diameter platform, a hollow airfoil and an inner diameter platform with a plurality of cooling tubes extending radially through the airfoil. The cooling tubes are open at the outer diameter end and closed with covers at the inner diameter end. The inner diameter platform is also cooled and includes a meterplate for a portion of the cooling passageway and includes an undercut to improve thermal deflections of the inner diameter platform.

Abstract: A stator blade for a gas turbine includes a blade airfoil extending in a longitudinal direction of the blade and is delimited by leading and trailing edges. The stator blade includes a shroud, whose inner side is exposed to hot gas which flows through the gas turbine. A hook-like fastening element for fastening the stator blade on a casing of the gas turbine projects outwards in the region of the trailing edge. The fastening element has a locating slot above the trailing edge for fixing a heat shield, which is connected to the shroud in the flow direction of the hot gas. A cavity is provided on the shroud between the locating slot for the heat shield and the trailing edge of the blade airfoil for reducing the thermal and mechanical stresses in the region of transition between the trailing edge and shroud.

Abstract: A turbomachine includes a housing, and a plurality of turbomachine nozzles. At least one of the plurality of turbomachine nozzles includes a body having a first surface and an opposing second surface that extend between a first end and a second end. At least one of the first and second surfaces includes a profile having an angle of between about 3° and about 8° between a first line passing through the first end and a first position located between about 28% and about 38% from the first end and a second line passing through the first position and a second position located between about 53% and about 63% from the first end, and an angle of between about 10° and about 20° between the second line and a third line passing through the second position and a third position located between about 74% and about 84% from the first end.

Abstract: A bidirectional centrifugal fan having a self-adjusting vane is disclosed. In one embodiment, the fan includes a driving disc having a pivot pin extending therefrom; a vane pivotably attached to the pivot pin, the vane including an aperture extending at least partially therethrough; and a trailing disc including a guide arm for receiving the aperture of the vane.

Abstract: A turbine nozzle includes: a hollow, airfoil-shaped turbine vane; and an arcuate first band disposed at a first end of the turbine vane, the first band having a flowpath face adjacent the turbine vane, and an opposed back face. The back face includes at least one open pocket, the at least one pocket defined in part by a bottom wall recessed from the back face, opposed ends of the bottom wall merging with the back face. The bottom wall is substantially free of interior corners.

Abstract: Second-stage nozzles have vanes including airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I with the X, Y and Z values commencing at the radially innermost aerodynamic section of the airfoil and then made relative to that section for the Z coordinate values. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the nozzle.

Abstract: Exemplary embodiments provide an information system adapted to focus a plurality of audio messages to a plurality of different locations. As a consumer passes a display, their attention is attracted based on the series of messages. Further, a consumer can obtain more and different information as they change their position. Exemplary embodiments may also display a different image to the consumer depending on the position of the consumer relative to the display. Further embodiments may utilize touch-screen technology to allow a user to select an icon from the screen and obtain further information either through additional messages, video, or both.

Abstract: A gas turbine engine is provided comprising an outer casing and a plurality of circumferentially positioned vane segments. The outer casing is provided with a circumferential casing slot. The plurality of circumferentially positioned vane segments are coupled to the outer casing. Each vane segment comprises at least one vane airfoil, a radially inner shroud coupled to a first end of the airfoil, a radially outer shroud coupled to a second end of the airfoil, and a strongback fixedly coupled to axially spaced-apart portions of the outer shroud such that a gap is provided between the strongback and the outer shroud. The strongback may comprise axially spaced-apart first and second end portions received in the casing slot.

Abstract: A two-stage, omnidirectional vertical axis wind turbine includes a first chamber in which a plurality of angularly spaced horns guide wind from any side of the wind turbine into a chamber to rotate a first rotor. Wind exiting the first rotor passes through a diverter to produce laminar flow. A fan attached to the shaft intercepts this laminar flow to add additional torque to the shaft.

Abstract: The monocrystalline turbomachine blade according to the invention that is cast and directionally solidified, is disclosed. The blade includes an airfoil with a leading edge, a pressure face, a suction face, a trailing edge, a skeleton and having a longitudinal axis, the faces and having a neck line, respectively a pressure face neck and a suction face neck relative to the adjacent blade in the turbomachine rotor of which it forms an element; an endpiece of the airfoil, such as a heel or a platform, having an airfoil end face, on the stream side, forming an angle with the axis ZZ; and a connection zone between the airfoil and the airfoil end face. The connection zone forms a fattening of the airfoil. The connection zone extends about the leading edge between a point P1 situated on the suction face of the airfoil upstream of the suction face neck and a point P3 situated on the pressure face of the airfoil upstream of the pressure face neck.

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y,Z? given in Table 1, in which the coordinate Z? is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: An actuation system includes a driving ring configured to rotate and having a groove on an internal face facing a central point of the driving ring; at least a linkage attached with a first end to an inside of the groove; and at least a lever arm attached to a second end of the at least a linkage. At least a portion of the at least a linkage stays inside the groove when the driving ring rotates.

Abstract: A turbine apparatus includes: A first nozzle comprising an array of first vanes each including a concave pressure side, a convex suction side, and leading and trailing edges; A rotor downstream from the first nozzle comprising a plurality of blades carried by a rotatable disk; and a second nozzle disposed downstream from the rotor comprising an array of second vanes each including a concave pressure side, a convex suction side, and leading and trailing edges; wherein the first and second vanes of the first and second nozzles are circumferentially clocked relative to each other such that, in a predetermined operating condition, wakes discharged from the first vanes are aligned in a circumferential direction with the leading edges of the second vanes, wherein a stacking axis of the first vanes is nonlinear. An inner band of the first nozzle is contoured in a non-axisymmetric shape.

Abstract: The pump assembly includes an impeller rotatably mounted in a pipe of the reactor having blades that generate a flow of fluid slurry through the pipe, a nosecone protruding in an upstream direction in front of the impeller blades that rotates along with the blades, and a plurality of stationary guide vanes connected to the housing upstream of the impeller blades and having free ends that are radially spaced apart from the nosecone. The leading edges of the stationary vanes are shaped to direct any stringy material into the radial gaps between the free ends of the stationary vanes and the rotating nose cone to be shredded and thus prevent fouling and clogging of the pump assembly when such stringy material is generated during the polymerization process.

Abstract: A method for optimizing the coating of a component using a coupon by introducing the coupon into a depression of the component, in which the geometry of the component is not altered as compared to a component without a coupon, is provided. A component having a depression into which a coupon is fastened is also provided.

Abstract: A technique for refurbishing nozzle diaphragm sections of a gas turbine replaces an eroded section of the nozzle diaphragm with a replacement part designed to engage a slot machined in the nozzle diaphragm. The replacement part is formed of a material with capable of sustained exposure to higher temperature than the original eroded section, and with a similar coefficient of expansion as the material used for manufacture the original nozzle diaphragm. The combination of the nozzle diaphragm and the replacement part conform to the original manufacturer's dimensional specifications for the nozzle diaphragm.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE A. X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches. The profile sections at the Z distances can be joined smoothly with one another to form a complete airfoil shape.

Abstract: An airfoil for a stator vane having an uncoated profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I is provided. The profile is carried only to three decimal places wherein Z is a distance from a platform on which the airfoil is mounted and X and Y are coordinates defining the profile at each distance Z from the platform.

Abstract: There is provided a turbocharger with a variable nozzle assembly having a plurality of cambered vanes positioned annularly around a turbine wheel, each vane (20) being pivotable around a pivot point (Pp) and being configured to have a leading edge (Ple) and a trailing edge (Pte) connected by an outer airfoil surface (2) and an inner airfoil surface (4), said outer airfoil surface (2) being substantially convex and said inner airfoil surface (4) having a convex section at the leading edge (Ple) which has a local extreme (Pex) of curvature and transitions into a concave section towards the trailing edge (Pte). The positions of the pivot point (Pp) and the local extreme (Pex) are set such that, even when the vanes are placed in a closed position, the exhaust gas stream exercises a positive torque on the vanes which tends to open the nozzle.

Abstract: A stator vane assembly for a compressor has a support structure which carries and is bounded by an annular stator vane structure. The stator vane structure comprises a central bore and a sleeve carried on the central bore. The sleeve is disposed between the support structure and bore of the annular stator vane structure. The annular stator vane structure is made from a non-metallic composite material and the sleeve is made from a first material. The coefficient of thermal expansion of the non metallic material is equal to or less than the co-efficient of thermal expansion of the first material.

Abstract: A gas turbine engine nozzle is described. The gas turbine engine nozzle includes at least one nozzle vane having a first end and a second end. The first end is coupled to an inner sidewall and the second end is coupled to an outer sidewall. The gas turbine engine nozzle also includes at least one stress relief pocket defined within at least one of the inner sidewall and the outer sidewall proximate to the at least one nozzle vane. The at least one stress relief pocket is configured to reduce stress on the proximate nozzle vane.

Abstract: A stator vane includes leading and trailing edges extending longitudinally between a root and a tip, the trailing edge having an exit swirl angle distribution such that the exit swirl angle has a maximum value at an intermediate radius location.

Abstract: A fan array fan section in an air handling system includes a plurality of fan units arranged in a fan array and positioned within an air handling compartment. One preferred embodiment may include an array controller program to operate the plurality of the fan units at peak efficiency. The plurality of fan units may be arranged in a true array configuration, a space pattern array configuration, a checker board array configuration, rows slightly offset configuration, columns slightly offset configuration, or a staggered array configuration.

Abstract: A vane for a turbine assembly of a turbocharger includes an airfoil that has a pair of flow surfaces disposed between a hub end and a shroud end and a leading edge and a trailing edge where the airfoil further includes a non-zero sweep angle, a non-zero lean angle, a non-zero twist angle or any two or more combinations thereof. Various other examples of devices, assemblies, systems, methods, etc., are also disclosed.

Abstract: To provide a turbocharger whose lifetime is long, which is capable of inexpensively and precisely performing nozzle vane angle synchronous control by using inexpensive and highly wear resistant components in a link mechanism for driving a rotation mechanism of nozzle vanes of an adjustable nozzle mechanism. As pins fixedly arrayed in a ring plate rotatably supported on the turbocharger, that set an angle of the nozzle vanes for controlling an amount of exhaust gas for rotating a turbine rotor, each pin in which a hard coating with a thickness of approximately 2 ?m which is formed of Al, Cr, Si, and N by physical vapor deposition is formed onto an austenitic stainless steel as a base material, is used.

Abstract: A guide device for a turbomachine, particularly a steam turbine, with a guide vane ring and a plurality of guide vanes which are fastened to the guide vane ring so as to be distributed along the circumference of the guide vane ring. The guide vane ring has a groove, and the guide vanes also have grooves which are aligned with the groove of the guide vane ring in the assembled state. The guide vanes are fixed in their radial position at the guide vane ring by means of a retaining ring which engages in the groove of the guide vane ring and in the grooves of the guide vanes.

Abstract: A straightener blade for a turbomachine, comprising a platform arranged at a radially external end of the blade and adapted to be engaged in a annular groove of a housing, and a threaded rod radially extending outwardly from the platform and adapted to pass through an orifice of the housing in order to receive a fixing nut, the upstream and downstream edges of the platform each comprising a protrusion extending in the plan of the platform in order to limit the rotation of the platform in the groove when the nut screwed on the threaded rod is tightened.

Abstract: An airfoil with turning flow comprises a first surface, a second surface, a leading edge and a trailing edge. The first and second surfaces extend from the leading edge to a trailing edge. A chord extends from a midpoint of the leading edge to a midpoint of the trailing edge, defining a duct angle of at least fifteen degrees with respect to a turbine axis. An axial length of the chord defines an aspect ratio of at most one with respect to an average radial span of the leading and trailing edges. In a region of the trailing edge, an axial angle is at most four degrees and a circumferential angle is at most ten degrees, where the circumferential angle is defined in a direction opposite the turning flow.

Abstract: A gas turbine engine component providing a gas flow passage is provided, which gas flow passage surrounds an inner passage surface, an outer passage surface forming an outer delimitation of said gas flow passage. In an aspect, the trace of the inner passage surface in a cross-section perpendicular to a central axis of the component presents at least one substantially straight portion. In another aspect, where a plurality of circumferentially spaced blades extends between the inner and outer passage surfaces, two portions of the trace, in a cross-section perpendicular to a central axis of the component, of the outer passage surface between two adjacent blades are substantially straight and oriented in an angle in relation to each other so as to form a concavity in the gas flow passage.