Abstract: Vibration induced compressor vane failure from tip leakage vortex bursting is eliminated or minimized. By securing the vane at opposite ends to inner and outer stationary casings, tip leakage is entirely avoided hence avoiding the mechanism for inducing vibration. By contouring the inner surface of the flow path to converge the flow in a downstream direction with a cantilevered compressor vane having a vane tip spaced from the inner casing surface, airflow lift off is precluded or minimized maintaining the flow attached to the flowpath surfaces with consequent avoidance of tip vortex induced vibration.

Abstract: A vane assembly for a gas turbine engine including a shroud with openings defined about a circumference of the shroud and vanes extending radially from the shroud with a vane tip in each opening. Isolating apparatus are disposed in each opening to isolate the vane tip from the shroud ring. Each vane tip is engaged with the shroud by retaining apparatus for restricting movement of the vane tip in an axial direction relative to the shroud.

Abstract: A conformal seal (20) for sealing air flow between a cooling airflow path and a hot gas flow path within a combustion turbine engine. The conformal seal (20) may be fitted within cooperating side slots of adjacent vane segments (10) within the combustion turbine engine. The conformal seal (20) may include an elongated metallic substrate (22, 40) forming an upper surface and a lower surface. A conformal coating (26, 44) may be deposited over one or both surfaces of the substrate (22, 40). The conformal coating (26, 44) may be deposited to a depth so that a point contact between the conformal coating (26, 44) and respective interior walls of the side slots wears the conformal coating (26, 44) to establish surface area contact there between. The surface area contact improves a sealing function between the conformal coating (26, 44) and the respective interior walls during operation of the combustion turbine engine.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a transition exit seal (302) adapted for fixed attachment to an exit rail (350) of a transition (109), and a stage-1 seal (340). The stage-1 seal (340) comprises an axial seal slot (348) comprising parallel walls (346, 347) adapted to slidingly engage a first male member (310) of the transition exit seal (302). The stage-1 seal (340) also comprises a radial seal slot (354) comprising parallel walls (351, 353) adapted to slidingly engage a plate (338) of a row 1 vane segment (330). The transition-to-turbine seal (300) thus provides both axial and radial freedom of movement. Materials and coatings for the transition exit seal (302) and the stage-1 seal (340) are selected to promote preferential wear of these components relative to wear of the surface of the row 1 vane segment (330). Wear of the exit rail (350) is reduced or eliminated due to the noted fixed attachment.

Abstract: A method for manufacturing a turbine stator vane includes coupling an airfoil to a vane stem wherein the airfoil includes a first sidewall and a second sidewall and wherein the first and second sidewalls are joined at a leading edge and at an axially-spaced trailing edge The method further includes forming an opening at least partially through the vane stem that is sized to receive a light probe therein.

Abstract: The first stage nozzles have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth Table I. The X and Y values are in inches and the Z value is in inches along the nozzle stacking axis coincident with a turbine radius. The X and Y distances may be scalable as a function of the same constant or number to provide a scaled up or scaled down airfoil section for the nozzle. The nominal airfoil given by the X, Y and Z distances lies within an envelope of±0.160 inches.

Abstract: Gas turbine systems having flexible chordal hinge seals are provided. According to an embodiment, a turbine system comprises: a nozzle segment comprising a stator vane extending between an inner band segment and an outer band segment; an inner support ring adjacent to the inner band segment; and an inner chordal hinge seal in operable communication with the nozzle segment, the inner chordal hinge seal comprising a flexible inner rail extending inwardly from the inner band segment, the inner rail having a projection for sealingly engaging the inner support ring.

Abstract: A turbine nozzle guide vane 10 with passages 28 leading from a hollow core 32 to respective seal strip slots 20, to deliver cooling air to abutment faces 18 on each end of the vane 10.

Abstract: Disclosed is an air intake for a turbocharger for an internal combustion engine, said air intake including a series of variable position guide vanes for directing the air such that it is swirling in a rotational sense on reaching the impeller of the compressor of the turbocharger. The variable position guide vanes are made comprised of sound absorbent material such as to achieve a substantial reduction in turbocharger noise at the air intake.

Abstract: The shape, of a vane for use in a variable nozzle assembly of a turbocharger was designed, including a step of defining a camberline curve and a thickness curve by means of two Bézier curves having a certain number of control points, and a step of applying computational fluid dynamics analysis and a Design of Experiments methodology for optimizing the setting of the control points to improve performance of the vane. As result, a distinct vane shape as shown in the drawings was obtained.

Abstract: The invention relates to a radial compressor having a suction-side opening and a pressure side. A swirl generator is placed upstream of the radial compressor. The swirl generator, has an inlet portion which is angularly configured relative to the outlet opening, the outlet opening facing the suction-side opening of the swirl generator. The swirl generator further has adjustable baffle plates arranged concentrically around the outlet opening.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in a TABLE 1. Wherein 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 being joined smoothly with one another to form a complete airfoil shape.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in a TABLE 1. Wherein 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 being joined smoothly with one another to form a complete airfoil shape.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in a TABLE 1. Wherein 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 being joined smoothly with one another to form a complete airfoil shape.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in a TABLE 1. Wherein 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 being joined smoothly with one another to form a complete airfoil shape.

Abstract: A method for repairing or replacing a mechanically retained vane is provided. The method comprises the steps of forming an oversized cavity in a support structure, inserting a flared end of a vane in the oversized cavity, and inserting a wedge for mechanically retaining the flared end of the vane in the oversized cavity.

Abstract: A turbine airfoil usable in a turbine engine and having at least one cooling system. At least a portion of the cooling system may be positioned in an endwall attached to the turbine airfoil. The endwall may include an endwall horseshoe cooling slot positioned in the first endwall proximate to the leading edge of the airfoil such that one end terminates proximate to the pressure side of the generally elongated hollow airfoil and a second end terminates proximate to the suction side of the generally elongated hollow airfoil. The endwall horseshoe cooling slot may include a plurality of film cooling holes angled to create a film cooling layer to reduce airfoil and endwall temperatures at the intersection between the leading edge of the airfoil and the endwall.

Abstract: A split shroud system for a gas turbine engine having a pair of annular-shaped shrouds that each have an inner pocket; to form a pair of inner pockets. Each of the pair of pockets having liner parts that form a circle. Liner parts of one of the pair of pockets facing liner parts of the other of the pair of pockets to form liner part pairs. One of each of said of pair of liner parts has a mutual abutting surface that forms a plurality of slots for accepting a plurality of vane inner trunnions.

Abstract: A turbine nozzle includes a row of vanes joined to radially outwardly inclined outer and inner bands. Each vane has camber and an acute twist angle for importing swirl in combustion gases discharged at trailing edges thereof. The trailing edges are tilted forwardly from the inner band to the outer band for increasing aerodynamic efficiency.

Abstract: A vane including a root, a tip and an airfoil portion extending between the root and the tip, the vane also including a button portion interconnecting the airfoil portion and the root. The button portion protrudes beyond a profile of the airfoil a distance less than the root. The button portion includes leading and trailing ends free of sharp edges. The leading and trailing ends respectively protrude beyond leading and trailing edges of the airfoil portion.

Abstract: A variable inlet guide vane assembly in a turbomachine in which the vanes can vary in angle to control the flow through into the turbomachine, each guide vane having a leading edge portion and a trailing edge portion, and where the entire vane can pivot or the trailing edge portion can pivot about the leading edge portion. The leading edge and trailing edge portions each have a magnetic field generating material located within the portion of the vane or one a surface of the vane. The vanes are offset in the shroud such that the trailing edge portion of a first vane is located closer to the leading edge portion of an adjacent vane than to the trailing edge portion of the adjacent vane. An electric current is passed through at least one of the trailing edge and leading edge portions to produce a magnetic field. The magnetic field causes the trailing edge portion to move toward the leading edge portion of the adjacent vane.

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: A cooling system is provided for automatic transmission torque converter assemblies having a housing shell and flex plate. The cooling system includes a housing cover having opposing exterior and interior surfaces that define a wall portion and a preferably circular base portion. The base portion is configured to attach to the flex plate, preferably by one or more stud members, to be rotated thereby. The wall portion extends from a periphery of the base portion and is configured to attach to the housing shell. The cooling system also includes a plurality of vane members positioned along the outer periphery of the base portion. The vane members are configured to increase convective dissipation of heat from the torque converter assembly to surrounding ambient air by increasing the heat transfer coefficient and surface area. Optimally, the vanes have a turbine fin configuration, but may alternatively have curved fin or a straight-fin configuration.

Abstract: Appropriate location of components such as vanes in gas turbine engines is important. These components are located between tangs held in interference engagement. There are four relatively rigid mounting points constituted by the tangs engaging respective opposed slots in a casing or a mounting ring. Only three point limitation with respect to displacement and rotation is adequate so provision of a fourth locator causes unnecessary constraint. A relatively resilient or compliant locator means it is possible to reduce the amount of constraint upon the component mounting whilst the other locators are adequate for positioning of the component.

Abstract: A turbocharger unit for a multicylinder internal combustion engine has at least two exhaust lines for the evacuation of exhaust gases from the combustion chamber of the engine and at least one inlet line for the supply of air to the combustion chamber. The turbocharger unit includes at least one turbine which cooperates with at least one compressor to extract energy from the exhaust flow of the engine and pressurize the inlet air of the engine. The turbine is provided with a worm-shaped housing having at least two flow paths which are defined by guide tongues in the housing and which conduct mutually separate exhaust flows via stator blades to at least one turbine wheel. In order to maintain the kinetic energy in the separate exhaust flows, each respective guide tongue meets a stator blade extending toward the tongue.

Abstract: A torque converter structure having a plurality of high efficiency secure weld free connections between blades and the shells of the pump and turbine that reduces flow between the blades and the shells. At least one blade within the torque converter is securely mechanically attached to the shell of the pump or turbine in the absence of welding by providing a tab on the blade that is bent on a portion that passes through a slit in the shell to secure the blade to the shell and by inserting a major part (greater than 50%) of a portion of an edge of the blade, that runs coextensively with a surface of the shell into a depression in that shell until the edge of the blade diverges from the shell. The invention also includes the unique blade and shell and a method for using them to make a torque converter of the invention.

Abstract: The first stage nozzles and buckets have nominal airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and R set forth in Tables I and II, respectively, and second stage nozzles and buckets have nominal airfoil profiles substantially in accordance with Cartesian coordinate values set forth in Tables III and IV, respectively. The X, Y and R values are in millimeters. R represents distances along a radius from an axis of rotation of the turbine. For each airfoil, X and Y are distances which, when connected by smooth continuous arcs, define airfoil profile sections in planes normal to the radius at each distance R. The profile sections at the R distances for each airfoil are joined smoothly with one another to form the airfoil shape. The airfoil shapes given by the X, Y and R distances lie within envelopes of ±4.064 millimeters in a direction normal to any airfoil surface location therealong.

Abstract: A turbomachine blade row is provided having a hub that includes a non-axisymmetric end wall modified by a transformation function. The blade row further includes a circumferential row of a plurality of airfoil members radially extending from the non-axisymmetric end wall of the hub and forming a plurality of sectoral passages therebetween. A radius of the non-axisymmetric end wall is determined by a transformation function including a plurality of geometric parameters defined by a user based on flow conditions. The plurality of geometric parameters provide for modification of the end wall in both an axial and a tangential direction to include a plurality of concave profiled regions and convex profiled regions.

Abstract: A device (10) for suspending gas channel elements, in particular for suspending guide blades or guide blade segments or gas channel segments, on a housing of a gas turbines is provided. The device comprises first plate-shaped elements (11, 12, 13, 14) and second plate-shaped elements (15, 16, 17), whereby the first plate-shaped elements (11, 12, 13, 14) and the second plate-shaped elements (15, 16, 17) are connected together by web-like elements (18) which extend in an essentially perpendicular manner in relation to the first and second elements and form a meandering or crenelated profile.

Abstract: The invention relates to a slotted bleed deflector for a gas turbine engine. The bleed deflector comprises an inlet portion for receiving bleed air from the engine and a body for distributing the bleed air into a fan bypass duct. The body includes a leading edge section, a trailing edge section and a flow compartment section where the flow compartment distributes the bleed air above an inner wall of the bypass duct.

Abstract: A gas turbine engine airfoil has a platform cooling scheme including an impingement hole for directing cooling air against an undersurface of the airfoil platform.

Abstract: A mounting arrangement (10) for a multivane segment (12) of ceramic matrix composite (CMC) composition positioned between outer and inner metallic rings (14, 16). Selected ones of the vanes (18a) of the multivane segment surround internal struts (24) joining the outer and inner rings. Spring members (26, 28) accommodate differential thermal growth between the multivane segment and the outer and inner rings, and a compliant material (30) seals against gas leakage around the segments.

Abstract: A guide vane for a gas turbine with a vane base body which is of single-piece design and comprises a profiled vane blade extending between a vane root and a cover plate and also the vane root formed integrally with the vane blade and the cover plate formed integrally with the vane blade, is intended, in a relatively simple way to be able to be matched to the individual conditions of use with especially little outlay on apparatus and logistics. For this purpose, according to the invention, a flow-routing body with an advance guide blade that is connected upstream of the vane blade as seen in the direction of flow of the working medium of the gas turbine is joined to the vane base body.

Abstract: A circumferential row of aerofoil members span an annular duct for carrying a flow of compressible fluid. The duct is centred on the axis of a turbomachine. Pressure and suction surfaces of neighbouring aerofoil members bound respective sectoral passages which receive the flow of compressible fluid. The row comprises 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.

Abstract: A nozzle guide vane for a gas turbine comprises at least one aerofoil having a leading edge section provided with at least one internal cavity for conveying cooling fluid through the said leading edge section, and a plate like wall section extending from the leading edge section to provide a continuous surface between the aerofoil leading edge section and the aerofoil trailing edge on the suction side of the aerofoil.

Abstract: The present invention provides a turbofan engine including leading edge erosion protection for composite stator vanes contained therein. The composite stator vanes are defined by a leading edge, a trailing edge, a concave surface and a convex surface. The vanes are generally formed of a graphite fiber and an aramid fiber braided preform that is laminated with an epoxy matrix resin or other high temperature matrix resin. A wire mesh screen is embedded during a molding process into an outermost surface of the leading edge of the epoxy matrix resin. A tapering erosion protection layer comprised of an erosion protective material is applied to the wire mesh screen at the leading edge of the composite stator vane an aids in protecting the leading edge of the stator vane from debris and contaminants, including operational fluids, dust, and sand that may be carried in an air stream passing therethrough the turbofan engine, as well as an erosive operational environment.

Abstract: A turbine housing of an exhaust gas turbocharger having an adjustable turbine geometry is provided having a carrier ring on which are mounted adjustable guide vanes for regulating the boost pressure and having a cover ring for the guide vanes, whereby the carrier ring and cover ring are spaced a distance apart by means of spacer elements to form a guide vane gap. A fastening screw passes through each spacer element, and screwed into the carrier ring and/or cover ring. Receiving openings for the head of the fastening screw are provided in the carrier ring or cover ring. At least one securing recess is provided in the carrier ring and/or cover ring on the outside circumference of the receiving opening, said securing recess being provided to accommodate material of the screw head to establish a form-fitting connection.

Abstract: An air intake with at least one inlet port area and one outlet port area, in which there is arranged at least one structure which may be flowed through and around, comprises an opening and modifies the clear cross-section only slightly, in particular serving as a fresh air/recirculated air housing for a motor vehicle heating, ventilation and/or air-conditioning system.

Abstract: A method for manufacturing a turbine engine component includes forming a component to include a first side and an opposite second side, positioning at least one capillary adjacent to an external surface of at least one of the first and second sides, and securing the at least one capillary to the component with at least one composite layer.

Abstract: An axial turbine includes a plurality of stages each comprising a plurality of stationary blades arranged in a row along the turbine circumferential direction and a plurality of moving blades in a row parallel to the stationary blades, each of the moving blade being disposed downstream of a respective one of the corresponding stationary blade in a flow direction of a working fluid so as to be opposed to the corresponding stationary blade. Herein, each of the stationary blades is formed so that the intersection line between the outer peripheral portion of the stationary blade constituting a stage having moving blades longer than moving blades in a preceding stage and a plane containing the central axis of the turbine, has a flow path constant diameter portion that includes at least an outlet outer peripheral portion of the stationary blade and that is parallel to the turbine central axis.

Abstract: A Mixer/Ejector Wind Turbine (“MEWT”) system is disclosed which routinely exceeds the efficiencies of prior wind turbines. In the preferred embodiment, Applicants' MEWT incorporates advanced flow mixing technology, single and multi-stage ejector technology, aircraft and propulsion aerodynamics and noise abatement technologies in a unique manner to fluid-dynamically improve the operational effectiveness and efficiency of wind turbines, so that its operating efficiency routinely exceeds the Betz limit. Applicants' preferred MEWT embodiment comprises: an aerodynamically contoured turbine shroud with an inlet; a ring of stator vanes; a ring of rotating blades (i.e., an impeller) in line with the stator vanes; and a mixer/ejector pump to increase the flow volume through the turbine while rapidly mixing the low energy turbine exit flow with high energy bypass wind flow.

Abstract: A steam turbine nozzle singlet having a blade or airfoil between inner and outer sidewalls is provided. The sidewalls include steps or flanges which are received in complementary recesses in the rings enabling axially short low heat input welds e.g., e-beam welds. These complementary steps and recesses mechanically interlock the singlet between the rings preventing displacement of the singlet in the event of weld failure. The low heat input welds minimize or eliminate distortion of the nozzle flow path. Additional features on the singlets, provide a datum for milling machines to form singlets of different sizes.

Abstract: A turbomachine blade with a profile skeleton line extending along a meridional flow line, the blade being radially divided into at least three zones (Z0, Z1, Z2) with profile skeleton lines of each zone (Z0, Z1, Z2) provided in each zone from the respective radially inner to the radially outer boundary to satisfy the equations: ? * = ? 1 - ? P ? 1 - ? 2 S * = s P S where P is any point of the profile skeleton line, ?1 is angle of inclination at blade leading edge, ?2 is angle of inclination at blade trailing edge, ?* is dimensionless, specific angle of total curvature, S* is dimensionless; specific extension, ?P is angle of tangent at any point P of profile skeleton line to central meridional flow line, sP is extension of profile skeleton line at any point P, and S is total extension of profile skeleton line.

Abstract: An axial flow turbine provided with a stage composed of a turbine nozzle and a turbine rotor blade arranged in an axial flow direction. Both end portions of a nozzle blade of the turbine nozzle are supported by a diaphragm inner ring and a diaphragm outer ring, and a flow passage is formed to have its diameter expanded from an upstream stage to a downstream stage. In such axial flow turbine, trailing edges at ends of the nozzle blade supported by the diaphragm inner ring and the diaphragm outer ring are curved as a curvature to an outlet side, and an intermediate portion between the trailing edges is formed to be straight.

Abstract: A vane assembly for a gas turbine engine including a shroud ring with openings and isolating arrangement isolating a vane extremity within each opening, the isolating arrangement cooperating to form a continuous gas path surface completely covering at least an axial portion of a surface of the shroud ring.

Abstract: A nozzle guide vane arrangement comprising a vane member having a platform portion extending forwards of a leading edge of the vane member, the platform portion having a first securing member and a second securing member to provide at least part of the means to retain the vane member in use, the arrangement characterised in that the first securing member and the second securing member are staggered with an axial displacement relative to each other upon the platform portion in the direction from the leading edge to a trailing edge of the vane member.

Abstract: A heat-dissipating device and a housing thereof. The housing includes a passage for guiding an air stream flowing from an opening to another opening, wherein an sidewall of the passage at least one of the opening sides extends radially outwards with a rotational axis of the heat-dissipating device or the passage so as to enlarge intake or discharge area for the air streams. Accordingly, the intake airflow rate may be greatly increased and the heat-dissipating efficiency of the heat-dissipating device may be greatly enhanced without changing assembling conditions with other elements.

Abstract: A stator for the second phase of a low-pressure turbine has a series of blades each defined by coordinates of a discreet combination of points, in a Cartesian reference system (X, Y, Z), wherein the axis (Z) is a radial axis intersecting the central axis of the turbine. The profile of each blade is identified by means of a series of closed intersection curves between the profile itself and planes (X, Y) lying at distances (Z) from the central axis. Each blade has an average throat angle defined by the cosine arc of the ratio between the average throat length at mid-height of the blade and the circumferential pitch evaluated at the radius of the average throat point; the average throat angle ranges from 57.8° to 60.8°.

Abstract: A turbine stage includes a stator nozzle cooperating with a row of turbine blades. The nozzle includes a row of vanes mounted between inner and outer bands, with the inner band adjoining the blades at inner platforms thereof at a rotary seal. The inner band includes a stator bullnose at the trailing edge thereof cooperating with the rotary seal for reducing aerodynamic losses thereat.

Abstract: A method for assembling a turbine nozzle for a gas turbine engine facilitates improving cooling efficiency of the turbine nozzle. The method includes providing a hollow doublet including a leading airfoil and a trailing airfoil coupled by at least one platform, wherein each airfoil includes a first sidewall and a second sidewall that extend between a respective leading and trailing edge. The method also includes inserting an insert into at least one of the airfoils, wherein the insert includes a first sidewall including a first plurality of cooling openings that extending therethrough, and a second sidewall including a second plurality of cooling openings extending therethrough, and wherein the first plurality of cooling openings facilitate more cooling of the airfoil than the second plurality of cooling openings.