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: The invention provides a method of fabricating a turbine stator casing, the method comprising the operations consisting in: between the walls of portions of a mold, forming a cavity of shape corresponding to the shape of the shroud of said casing, securing soluble cores to at least one of said mold portions, said cores being held at a distance from the wall of said mold portion and representing empty spaces that are to be formed inside said shroud; putting soluble inserts into place between the cores to represent flow paths between said empty spaces; filling said cavity with a metal alloy powder; sintering said powder by hot isostatic pressing; eliminating the cores and the inserts by dissolving them; and extracting the shroud as molded in this way. The invention is applicable to fabricating a turbine stator casing for an airplane turbojet.

Abstract: A high-lift airfoil embodying a combination of loading conventions in a single design specifically to reduce and control total pressure losses that occur in the flow channels between airfoils employed in turbomachinery applications is disclosed herein. The mixed-loading high-lift airfoil designs embody and exhibit the best total profile and secondary loss characteristics possessed by both aft-loaded airfoil and front-loaded airfoil conventions through controlling the development and interaction of boundary layers forming along the surfaces of the airfoils and the endwalls in such applications. The mixed-loading high-lift airfoil may be utilized in both rotating and non-rotating turbomachinery applications.

Abstract: A process of designing a transonic airfoil that results in an exit deviation angle that is optimized to minimize downstream shock induced flowfield disturbances. Specifically, a design process of a transonic airfoil minimizes pitchwise variation in downstream static pressure that includes determining an airfoil exit deviation angle and downstream turning for reduced shock from a trailing edge of the transonic airfoil.

Abstract: A system and method for orienting first and second reaction turbines relative to an axis of rotation responds to a fluid flow having a flow orientation axis. The fluid flow is received in a casement. The casement has first and second endplates situated parallel and spaced apart along the axis of rotation. A first half of the fluid flow drives the first reaction turbine to rotate with a first spin orientation in a plane perpendicular to the axis of rotation to produce a first torque about the axis of rotation. The second half of the fluid flow drives the second reaction turbine, offset from the casement plane relative to the first reaction turbine, with a second spin orientation opposite the first spin orientation to produce a second torque about the axis of rotation. The casement is oriented about the axis of rotation in response to the first and second torques.

Abstract: An inlet guide vane having a nominal profile substantially in accordance with Cartesian coordinate values of x, y, and z set forth in TABLE 1 reduces vibration of engine components at various operating conditions. A scaling factor can be applied to the values to make the airfoil larger or smaller. The TABLE 1 x and y values are distances in inches within a tolerance which, when connected by smooth curves, define airfoil profile sections at each distance z in inches, the profile sections being joined smoothly to form a complete airfoil shape. The tolerance in an embodiment is up to about 0.16 inches.

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 turbine vane assembly includes an airfoil extending between an inner shroud and an outer shroud. The airfoil can include a substructure having an outer peripheral surface. At least a portion of the outer peripheral surface is covered by an external skin. The external skin can be made of a high temperature capable material, such as oxide dispersion strengthened alloys, intermetallic alloys, ceramic matrix composites or refractory alloys. The external skin can be formed, and the airfoil can be subsequently bi-cast around or onto the skin. The skin and the substructure can be attached by a plurality of attachment members extending between the skin and the substructure. The skin can be spaced from the outer peripheral surface of the substructure such that a cavity is formed therebetween. Coolant can be supplied to the cavity. Skins can also be applied to the gas path faces of the inner and outer shrouds.

Abstract: A method for repairing or replacing a mechanically retained vane is provided. The method comprises the steps of forming an oversized cavity in an outer base, 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. The wedge has a first surface with a constant pitch angle to match the outer base cavity and a second surface with a variable pitch angle to match the vane tip dovetail.

Abstract: Platforms with curved side edges and gas turbine engine systems involving such platforms are provided. In this regard, a representative airfoil assembly for a gas turbine engine includes: a platform having a gas path side, a non-gas path side, a leading edge, a trailing edge, a first side edge extending between the leading edge and the trailing edge and exhibiting a first curve along a length thereof, and a second side edge extending between the leading edge and the trailing edge and exhibiting a second curve along a length thereof; and an airfoil extending from the gas path side of the platform; the platform and the airfoil exhibiting a unitary construction such that a continuous exterior surface blends from the airfoil to the platform.

Abstract: The present invention provides an airfoil a second stage nozzle guide vane having an external surface with first and second sides. The external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge of the airfoil. The external surface includes a contour substantially defined by Table 1 as listed in the specification.

Abstract: A gas turbine stator includes a vane assembly, and a platform that is configured to receive the vane assembly. The vane assembly includes a vane and a cap. The vane extends radially from the cap. The vane includes a pressure side shell and a separate suction side shell.

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 flow-guiding member unit used for assembling a flow-guiding member comprising pluralities of vanes, an outer platform and an inner platform, the unit comprising an integral body member comprising one vane portion, an outer platform portion, and an inner platform portion, a first elastic member attached to a step of the outer platform portion, and a second elastic member attached to a step of the inner platform portion, the body member being formed by the thermal compression of sheet-molding compounds or fiber-reinforced resin pellets, and the first and second elastic members being made of a thermoplastic elastomer.

Abstract: A rotary machine includes a rotor, a stationary machine casing extending around the rotor, and a bling assembly extending between the casing and the rotor. The machine also includes at least one rotor tip seal assembly and at least one shaft seal assembly. The seal assemblies have a groove configured to receive at least one seal ring band. A method of assembling a rotary machine is also provided. The method includes fabricating the bling assembly by providing two identical members comprising a mating surface and having a semi-circular profile. The method also includes coupling the two members together at their mating surfaces such that a circular ring is formed and such that the mating surfaces define a horizontal joint. The method further includes machining concentric, circular and annular radially inner and outer and airfoil portions within predetermined radial portions of the bling assembly.

Abstract: What is described is a turbine guide vane having a guide vane leaf and with a radially outer platform connected to the guide vane leaf and having a platform surface which faces radially away from the guide vane leaf and on which a connecting structure is provided for fastening the guide vane to a supporting structure which has side wall portions which project radially beyond the platform surface and delimit an inner cavity and on which is provided at least in portions a joining contour which can be inserted into a reception shape countercontoured within the supporting structure. At least one seal is provided between the connecting structure and the supporting structure.

Abstract: A wind power engine including a stator having stator blades and a rotor having rotor blades. The rotor is positioned inside the hollow area of the stator and is rotationally movable with respect to the stator about an axis of rotation. The stator blades are positioned in such a way that an air flow coming from the outside of the stator is oriented towards the rotor and the rotor blades. The oriented airflow allows the rotor blades to drive the rotor about the axis of rotation. At least certain rotor blades have a concave-shaped profile on one of their faces.

Abstract: A nozzle for radial inflow turbines having an outer surface having a concave shape, and an inner surface having a non-concave, i.e. straight or convex, shape. Also a leading edge may have a hole that is capable of receiving a fastener to secure the nozzle vane to an assembly housing.

Abstract: A compressor vane having a root, a tip and an airfoil portion extending between the root and the tip is disclosed. A grommet is used in conjunction with the vane and has a lip sealingly surrounding a portion of the vane root and a circumferential portion protruding around a perimeter of the lip in engagement with a surface of a platform of the vane root. The circumferential portion has a profile similar to that of the airfoil portion.

Abstract: A method of forming a blade assembly for a gas turbine engine, the method comprising mounting turbine blades on a rotor in a number of radial stages. Sacrificial supports are placed on the blades, and nozzle guide vanes are located on the supports. The rotor is introduced into a casing, and the guide vanes are mounted to the casing.

Abstract: A Mixer/Ejector Wind Turbine (“MEWT”) system is disclosed which routinely exceeds the efficiencies of prior wind turbines. Unique ejector concepts are used to fluid-dynamically improve many operational characteristics of conventional wind turbines for potential power generation improvements of 50% and above. 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. The MEWT can produce three or more time the power of its un-shrouded counterparts for the same frontal area, and can increase the productivity of wind farms by a factor of two or more. The same MEWT is safer and quieter providing improved wind turbine options for populated areas.

Abstract: A variable geometry inlet guide vane for a gas turbine aircraft engine includes an aerodynamic shell for turning inlet flow to a turbine or compressor and an internal spar spaced from the airfoil shell by an air gap. A number of actuation mechanisms grounded to the spar and connected to the inner surface of the aerodynamic shell adjust the shape of the shell in response to varying operating conditions of the engine, imbalanced aerodynamic loading of the shell or vibration or other transient loads on the shell.

Abstract: A vane for a vane assembly of a gas turbine engine includes a vane root connected to the airfoil portion opposite a tip thereof. The vane root includes a platform and a button portion interconnecting the platform and the airfoil position. The button portion has relatively blunt leading and trailing ends which respectively protrude beyond leading and trailing edges of the airfoil portion.

Abstract: A stator, including a blade assembly with an inner circumferential surface and a plurality of pre-formed grooves in the inner circumferential surface and an outer race with an outer circumferential surface and a plurality of protrusions at least partially engaged with the plurality of pre-formed grooves. The plurality of protrusions is frictionally engaged with the plurality of pre-formed grooves. In one embodiment, the inner circumferential surface includes an area circumferentially disposed between first and second protrusions from the plurality of protrusions and a portion of the area is displaced radially inward by the engagement of the pluralities of grooves and protrusions.

Abstract: A turbine blade cascade structure includes a plurality of blades arranged in series in a circumferential direction on a wall surface, in which a corner portion defined by the wall surface and a front edge portion of each of blade bodies supported by the wall surface, to which a working fluid flows, includes a cover portion (fillet) that extends toward an upstream side of a flow of the working fluid. The turbine blade cascade structure is capable of reducing the secondary flow loss of the secondary flow in spite of the fluctuation of an incident angle of the working fluid flowing to the front edge portion of the blade body.

Abstract: Embodiments for a compressor stator vane assembly in a gas turbine engine are disclosed. In an embodiment of the present invention a stator vane assembly is provided having a plurality of vanes each with an attachment and channels machined into the forward and aft walls of the attachment. A forward hook ring segment is pressfit into the channel in the forward wall of the attachment and an aft hook ring is pressfit into the channel in the aft wall of the attachment. The hook ring segments join a plurality of vanes together so as to provide a uniform engagement of mounting slots in the compressor case. Such an arrangement increases the contact area between the hook rings and the compressor case such that damping of individual vane vibrations are improved and operating stresses are reduced.

Abstract: A flow direction system (200) for a turbocharger (1) or other fluid impeller device is provided that can reduce swirl or turbulence of exhaust gases being supplied to a turbine rotor (4). The system (200) uses one or more ribs (250) or fins that extend from an inner surface (110) of an inlet (100) of the turbine housing (2) and that can be integrally formed with the inlet (100). The ribs (250) can be diametrically opposed along the inner surface (110) and can extend less than half way across the inlet (100). The ribs (250) can have chamfered leading or trailing edges to reduce drag.

Abstract: A replacement fan exit guide vane assembly for installation between an inner case and an outer case of a gas turbine engine includes an airfoil structure having an outer end and an inner end, a wedge structure connected to the outer end of the airfoil structure, a base plate positioned between the outer end of the airfoil structure and the outer case, a conformable material, and a slip collar having an opening therethrough and a recess formed along the opening. The airfoil structure extends through the opening, and the recess accepts the wedge structure to create an interference fit between the slip collar and the base plate. The conformable material is disposed along the opening and the recess in the slip collar, and also along the baseplate, in order to provide dampening relative to the airfoil structure and the wedge structure.

Abstract: A fluid flow machine includes at least one rotor blade row designed for high work coefficients, which—in its center section and/or averaged over the blade height—provides for an essentially swirl-free relative outflow. The respective rotor, in its trailing edge area, is essentially orientated in meridional direction, which essentially corresponds to the meridional flow direction.

Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.

Abstract: A two-stage high pressure turbine includes a first stage vane 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 turbine vane array (10) for a combustion assembly (8) in a combustion turbine engine. The vane array (10) includes a plurality of stationary vane assemblies (12), each vane assembly (12) including at least one airfoil (24) and inner and outer shroud segments (26, 28) attached to opposing ends of the airfoil (24). The inner and outer shroud segments (26, 28) each include an inner face (34, 36) facing toward a gas path (13) extending through the vane array (10). A removable cover structure may be provided on the inner faces (34, 36) of the inner and outer shroud segments (26, 28). The cover structure may include removably attached insert elements (72, 72?) that are positioned on the inner faces (34, 36) extending between upstream edges (44) and downstream edges (46) of the shroud segments (26, 28) and extending between adjacent airfoils (24).

Abstract: A ceramic matrix composite nozzle assembly. The ceramic matrix composite nozzle assembly may include a ceramic matrix composite vane, a number of metallic components positioned about the ceramic matrix composite vane, and a number of metallic seals positioned between the ceramic matrix composite vane and one or more of the metallic components.

Abstract: A turbomachine system comprises a first turbocharger comprising an exhaust gas flow first turbine for location in an exhaust path and a first compressor driven by said first turbine; a turbomachine for location in the exhaust path upstream or downstream of said first turbocharger and comprising an exhaust gas flow second turbine and a second compressor driven by said second turbine. The first turbine has an outlet that is in fluid communication with an inlet of the second turbine. One of said first and second turbines is a radial outflow turbine. The arrangement provides for a relatively compact package. The radial outflow turbine may have a particular structure in which there is provided a deflector member at or near its inlet for directing the gas outwards, a stator for introducing swirl and a downstream turbine rotor. A shroud is fixed to blades of the turbine rotor to prevent leakage and to provide additional structural rigidity.

Abstract: A gas-turbine axial compressor has a casing 3 and a rotatable shaft 5, which form an annular duct, in which at least one stator 2 and one rotor 4 are arranged. A shroud 21 is arranged at a free end of the stator 2 vanes, which extends over part of the axial length of the stator 2. A further part of the axial length of the stator 2, at which no shroud 21 is arranged, radially adjoins a rotor platform 28 formed by a rotor hub 8.

Abstract: A method is provided for attaching a shim to at least one of a base of a stator vane unit and a ring segment of a ring segment assembly in a gas turbine compressor. The method includes providing a shim shaped to conform to an engaging face of the base of the stator vane unit or the segment ring of the segment ring assembly, with a predetermined thickness dimension; boring at least one hole, normal to the face of the shim and extending fully through the thickness dimension of the shim; positioning the shim on the engaging face of the at least one of the base of the stator vane unit and the segment ring of the segment ring assembly; and welding the shim to the engaging face of the at least one of the base of the stator vane unit and the ring segment of the stator ring assembly.

Abstract: A passive internal control system for rotor blade tip clearance of a high-pressure turbine is based on inner rings that expand in a radial direction under the influence of heat and radially adjust the clearance delimited by liner segments (9) on the casing side. It comprises a radially expandable U-shaped downstream inner ring (10) mounted to inner platforms (7b) of guide vanes (7) on a side where the rotor does not have a static bearing, said ring providing expansion compensation in the axial and peripheral directions and forming a torsion box (8) with struts (13, 14) that absorbs rolling and tilting moments.

Abstract: Aspects of the invention relate to a turbine vane in which the inner and outer platforms are located substantially entirely on either the pressure side or the suction side of the airfoil. When a plurality of such vanes are installed in the turbine, a seam is formed by the circumferential end of the inner and outer platforms and a portion of the airfoil of a neighboring vane. During engine operation, a high pressure coolant is supplied to at least one of the platforms. The coolant can leak through the seam. Because the seam is located proximate the airfoil, the coolant leakage through the seam can be productively used to cool the transition region between the vane platforms and the airfoil. In addition to such cooling benefits, aspects of the invention can result in a potential increase in engine efficiency as well as component life.

Abstract: A turbine stage includes a stator nozzle having a row of vanes mounted between inner and outer bands. The inner band terminates in an ovate ledge converging aft from the vanes with radially outer and inner convex surfaces joined at a convex apex. The ovate ledge reduces aerodynamic losses at the rotary seal with a row of following turbine rotor blades.

Abstract: A centrifugal compressor assembly for compressing refrigerant in a 250-ton capacity or larger chiller system comprising a motor, preferably a compact, high energy density motor or permanent magnet motor, for driving a shaft at a range of sustained operating speeds under the control of a variable speed drive. Another embodiment of the centrifugal compressor assembly comprises a mixed flow impeller and a vaneless diffuser sized such that a final stage compressor operates with an optimal specific speed range for targeted combinations of head and capacity, while a non-final stage compressor operates above the optimum specific speed of the final stage compressor.

Abstract: An inlet guide vane is installed at a suction port into which a fluid is sucked by the rotation of an impeller so as to be rotatable about an axis, and is comprised of a rod-like shaft supported by an inner peripheral surface of the suction port so as to be rotatable about the axis, and a plate-like vane body connected to the shaft and provided so as to protrude from the inner peripheral surface of the suction port. The vane body includes a tapered portion formed such that both side surfaces approach each other toward an outer edge of the vane body in its width direction and an outer edge thereof at a tip in the direction of the axis, and a parallel portion arranged on the axis and formed such that both the side surfaces are parallel to each other along the direction of the axis.

Abstract: A single stage high pressure turbine vane includes an airfoil having 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 gas turbine includes a nozzle vane and a sealing unit engaged with the nozzle vane inside a turbine supplied with combustion gases produced by mixing and burning air for combustion and fuel. The nozzle vane and the sealing unit are disposed in a channel of the downward flowing combustion gases on the outlet side of a gas path. A plurality of engagement portions between the sealing unit and the nozzle vane are provided successively from the upstream side toward the downstream side in a direction of flow of the combustion gases, and a downstream one of the plurality of engagement portions has a contact interface formed in a direction across a turbine rotary shaft. A reduction in the thermal efficiency of the gas turbine can be suppressed.

Abstract: A device for varying the section of the throat of a turbine nozzle, the nozzle being formed by a plurality of stationary vanes extending radially between outer and inner annular platforms and spaced apart from one another in order to define a throat presenting a minimum flow section. The device comprises an annular element having a coefficient of expansion that is less than that of the platforms of the nozzle, the annular element being secured to the outer platform and being capable of taking up two positions: one position corresponding to no expansion of the platforms, in which the element provides continuity for the profile of the flow passage; and another position corresponding to the platforms having expanded, in which position the element projects into the flow passage so as to reduce its section.

Abstract: A nozzle segment is disclosed that includes a first platform configured to be connected with a casing of a gas turbine engine, a second platform configured to be connected with the casing, the second platform being disposed apart from the first platform in a radial direction of the gas turbine engine, and a plurality of blades disposed between the first and second platforms and connected thereto, a portion of an inside surface of the first platform or the second platform being scalloped so as to increase a throat area of the nozzle segment in order to increase the nozzle segment's mass flow rate handling capability.

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 turbine nozzle includes a row of vanes extending radially in span between inner and outer bands. The vanes include opposite pressure and suction sidewalls and opposite leading and trailing edges. Each vane includes an inner pattern of inner cooling holes and an outer pattern of outer cooling holes distributed along the leading edge. The inner and outer holes diverge toward the corresponding inner and outer bands to preferentially discharge cooling air.

Abstract: A nonlinearly stacked low noise turbofan stator vane having a characteristic curve that is characterized by a nonlinear sweep and a nonlinear lean is provided. The stator is in an axial fan or compressor turbomachinery stage that is comprised of a collection of vanes whose highly three-dimensional shape is selected to reduce rotor-stator and rotor-strut interaction noise while maintaining the aerodynamic and mechanical performance of the vane.

Abstract: An axial turbine includes a plurality of stages, each of the plurality of stages having stationary blades adjacent to each other along the turbine circumferential direction and corresponding moving blades adjacent to each other along the circumferential direction, each of the moving blade being located downstream of a corresponding one of the stationary blades along a flow direction of a working fluid, so as to be opposed to the corresponding stationary blade. In this axial turbine, each of the stationary blades is formed so that the intersection line between the outer periphery 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 periphery of the stationary blade and that is parallel to the central axis of the turbine.

Abstract: A tip turbine engine (30) includes a low pressure compressor (22) having a plurality of inlet guide vanes (55) that are mounted at an inlet to the compressor case (50). Each inlet guide vane (55) includes at least one fluid outlet (56) proximate a trailing edge of the inlet guide vane (55), such that fluid flow through the fluid outlet (56) modulates and controls the air flow into the compressor (22). A supply of pressurized fluid may be supplied from compressed air from the compressor (22).