Abstract: In an electronic system, a method for operating a cooling fan comprises rotating an impeller about a rotational axis and detecting fan failure. The impeller is spatially expanded in response to the detected fan failure whereby airflow through the failed fan is blocked.

Abstract: Gas flow is redirected by a feature disposed on a trailing edge of at least one segment of a peripheral gas path defining surface to improve alignment with a downstream portion of the gas path.

Abstract: A seal arrangement for reducing the seal gaps within a rotary flow machine, preferably an axial turbomachine, has rotor blades and guide vanes, which are respectively arranged in at least one rotor blade row and guide vane row and have respective blade/vane roots (2,3) which protrude into fastening contours within the rotor blade and guide vane rows. A sealing element (4) of plastically deformable material is provided between at least two adjacent blade/vane roots (2,3) along a rotor blade row or guide vane row or between a blade/vane root (2,3) of a rotor blade or guide vane and a rotary flow machine component directly adjoining the blade/vane root.

Abstract: A turbine nozzle includes a hollow vane joined to opposite inner and outer bands. The inner band includes a forward lip under the leading edge of the vane, an aft lip under the trailing edge of the vane, and a mounting flange spaced therebetween. An impingement bore extends obliquely and aft from the base of the mounting flange under a corner of the inner band toward the aft lip for impingement cooling thereof.

Abstract: A turboshaft engine, such as for example a turbine, including a segmented seal between two longitudinally adjacent elements such as, for example, a casing and a ring of sectors of vanes. According to one embodiment, a seal is interposed between a casing and a ring of sectors of vanes and includes a plurality of flexible segments having a straight deformation zone and a straight fluid-tight bearing zone.

Abstract: A sky turbine, having impeller blades with a plurality of panels that can telescope inward and outward. The pitch of the panels can be adjusted, and the pitch of each panel on a blade, passing from the innermost panel to the outermost panel, differs from the pitch of the previous panel by a certain adjustable amount. The sky turbine is preferably mounted above a city at a sufficiently high elevation that air current will be generally faster than just above the ground. In the preferred embodiments, a wing with two pods is pivotally mounted on a tower. Each pod has front and rear shafts, from which four blades extend. Servomotors cause the wing to turn so that its center of symmetry is parallel to the direction the sky is moving, cause the pitch of the panels of the blades to change, and cause the panels of the blades to telescope.

Abstract: A turbine nozzle includes a vane joined to opposite inner and outer bands. The inner band includes a pocket adjoining a flange which recesses in part the inner surface of the band. The pocket undercuts the flange to form a notch at the base thereof.

Abstract: A nozzle vane for a rotary disk of a turbomachine, the vane comprising pressure side and suction side surfaces extending radially between a base and a tip and longitudinally between a leading edge and a trailing edge, with a plurality of vane sections having centers of gravity aligned along a stacking axis, the vane presenting a lower portion, an intermediate portion, and an upper portion, the lower portion extending radially between the base of the vane and a lower limit of the intermediate portion, and the upper portion extending radially between an upper limit of the intermediate portion and the tip of the vane, the stacking axis presenting, in the lower and upper portions, a tangential component that is substantially radial, and in the intermediate portion, a tangential component with two slopes.

Abstract: An easily producible turbine nozzle for a gas turbine engine is provided, which is capable of preventing flutter of the turbine nozzle during operation of the gas turbine engine. The turbine nozzle comprises airfoils stacked along the stacking axis. The high curvature portions on suction surface in airfoil section successively formed along the stacking axis of the airfoil describe a parabola line that curves toward the pressure side of the airfoil when seen from the front or rear of the turbine nozzle. The high curvature portions on suction surface in airfoil section curve most at the center along the stacking axis of the airfoil from a straight line that connects with a first intersection between the parabola line and an inner band in the turbine nozzle and with a second intersection between the parabola line and an outer band in the turbine nozzle. The maximum curvature falls within a range from 0.02 to 0.03-fold of the stacking axis of the airfoil.

Abstract: A device for injecting cooling air into a turbo-machine turbine rotor includes a plurality of injectors distributed regularly around a longitudinal axis of the turbomachine and mounted between an inner shroud and an outer shroud. Each injector has a bimetallic structure with a first metal material forming a major portion of the structure of the injector and having a first coefficient of thermal expansion, and a second metal material forming a complementary portion of the structure in the vicinity of the suction side wall meeting the trailing edge of the injector, and having a second coefficient of thermal expansion that is smaller than the first.

Abstract: A turbine nozzle support structure supports a ceramic turbine nozzle provided with a flange and disposed at the inlet of a turbine on a nozzle support. Force-multiplying mechanism presses the flange of the turbine nozzle against a nozzle support to support the turbine nozzle on the nozzle support. The force-multiplying mechanisms includes a pressing member having a spring holding part, a fulcrum part and a pressing part, a spring for applying resilient force to the spring holding part, and a fulcrum supporting member having a bearing part for supporting the fulcrum part.

Abstract: A method enables a gas turbine engine to be operated. The method comprises supplying cooling fluid into a manifold ring that includes a plurality of distribution ports defined by a sidewall connected by a radially inner wall, channeling the cooling fluid circumferentially through the manifold ring and through at least one distribution port that is defined by a wall that extends arcuately across at least one turbine nozzle, and discharging cooling fluid from the distribution ports radially inwardly towards the at least one turbine nozzle positioned radially inward from the manifold ring.

Abstract: A sealing arrangement that can be employed to maintain a seal in any suitable pressure vessel including but not limited to a gas turbine engine. The sealing arrangement is forg sealing a leakage gap between at least two relatively moveable parts that are adjacent to each other in a flow path between a region of high fluid pressure and a region of low fluid pressure. The sealing arrangement comprises at least two resilient sealing strips, each strip having a portion formed along at least part of its width to locate in a groove provided along adjacent faces of the relatively moveable parts. The remaining portion of each of the sealing strips has a substantially flat surface which, in operation, abuts a corresponding flat surface of an adjacent sealing strip, thereby forming a seal.

Abstract: A vane assembly for a gas turbine is described wherein the vane assembly comprises a first vane and a second vane connected together by a plurality of flanges, at least one fastener, and at least one spring plate. The fastener and hole diameters in the respective flanges are sized such that the first vane and second vane are essentially pinned together along their inner flanges and allowed to adjust due to thermal growth along their outer flanges, while maintaining a constant seal along both inner and outer platform edges. The thermal growth along the outer flanges is made possible by oversized flange holes relative to the diameter of the fastener.

Abstract: A method of assembling a gas turbine engine includes providing a plurality of stator vane sectors that each include an equal number of stator vanes that are circumferentially-spaced such that a first circumferential spacing is defined between each pair of adjacent stator vanes within the sector, and coupling the plurality of stator vane sectors together to form a stator vane assembly such that a second circumferential spacing is defined between each pair of adjacent stator vanes coupled to adjacent sectors, wherein the second circumferential spacing is different from the first circumferential spacing.

Abstract: Nozzle segments are secured to a retention ring against circumferential rotation by anti-rotation pins extending generally axially between the outer bands of the nozzle segments and the retention ring. Retention plate segments overlie the ends of the pins, preventing axial removal thereof. To remove a selected nozzle segment, inner diameter retention plate segments and selected retention plate segments are removed, the latter exposing the ends of the pins for axial withdrawal. Upon removal of a predetermined number of pins, the nozzle segments adjacent the selected segment are displaced away from the latter segment to open a gap between the selected segment and adjacent segments whereby the selected segment can be removed in an axial direction.

Abstract: A turbine flowpath structure includes a first generally U-shaped flowpath channel having a first leg having a first inwardly facing side shaped as a turbine blade suction-side airfoil surface, a second leg having a second inwardly facing side shaped as a turbine blade pressure-side airfoil surface, and a web connecting the first leg and the second leg, wherein the web has an inwardly facing inner flowpath surface. There is typically provided a spar that engages at least one of the first leg and the second leg. In one application, there is a plurality of the generally U-shaped flowpath channels arranged around the periphery of a turbine disk, and a plurality of spars positioned between each pair of adjacent generally U-shaped flowpath channels to anchor them to the turbine disk.

Abstract: A stator blade for a hydrokinetic torque converter, the blade profile having geometric characteristics that improve torque flow efficiency in the converter torque flow circuit, whereby the torque converter efficiency at high speed ratios is improved without significant adverse effects on the converter torque ratio at stall and at low speed ratios.

Abstract: A vane assembly includes a vane having a main axis, a first passageway within the vane substantially parallel to the main axis, and a second passageway aligned with the first passageway. The vane has a window portion between the first passageway and the second passageway. The vane assembly also includes a sensor inserted in the first passageway and having a portion held in place in the second passageway, wherein a portion of the sensor is exposed in the window portion of the vane.

Abstract: A collar seal is configured for a turbine nozzle vane. The seal includes a retainer having a circumferential airfoil contour conforming with the airfoil contour of the vane. A flexible leaf is surrounded by the retainer and fixedly joined thereto. A woven sheath encases the leaf and is fixedly joined to the retainer. In an exemplary embodiment, the collar seal surrounds one end of a ceramic turbine vane mounted in a metal supporting band.

Abstract: A heat-dissipating device and a housing thereof. The heat-dissipating device includes an impeller and a housing. The impeller includes a hub and a plurality of rotor blades disposed around the hub. The housing includes an outer frame, an air-guiding part, and a plurality of stator blades. The air-guiding part is formed on at least one side of the outer frame. The stator blades are disposed in the outer frame corresponding to the air-guiding part. The stator blades guide the airflow through the blades, and greatly increase the static pressure of the airflow discharged from the heat-dissipating device.

Abstract: The invention provides a device for holding strip sealing gaskets on a turbomachine nozzle, the nozzle being constituted by a plurality of sectors having guide vanes which extend between inner and outer platforms elements, at least one of said platform elements having a sealing strip (30) disposed radially in a space between said platform element and an adjacent structural element in order to provide fluid tightness between said elements, the corresponding strips of two adjacent sectors touching, and their joint being covered by a joint cover (31) disposed between said strips, and by a holding member for holding said strips in the sealing position.

Abstract: A low-pressure compressor exit stator assembly for a gas turbine includes a multiple of vanes mounted between a full hoop outer shroud and an inner shroud. The outer shroud and inner shroud are mounted to the intermediate case structure. A seal is located between the outer shroud and a bleed duct structure to accommodate thermal growth mismatch.

Abstract: A turbine disk cooling device fed with cooling air from an orifice through an annular support platform for a fixed vane of a low-pressure turbine is disposed between an upstream flange and a downstream flange of the support platform. The device includes upstream and downstream annular plates longitudinally defining an annular cavity, a sealing device extending longitudinally between the upstream and downstream plates so as to close the cooling air cavity, an element to hold the upstream and downstream plates against upstream and downstream flanges of the support platform, and a plurality of holes for ejecting cooling air towards the turbine disks.

Abstract: A method of tuning a compressor stator blade having a base portion and an airfoil portion to achieve a desired natural frequency, includes a) identifying the natural frequency of the compressor stator blade; b) determining a different target natural frequency for the compressor stator blade; and c) removing material from the base portion of the compressor stator blade in an amount and in a configuration that achieves the target natural frequency. A frequency-tuned compressor stator blade includes an airfoil portion and a base portion, the base portion having a substantially solid rectangular shape; and a groove cut across a width dimension of the base portion, the groove having dimensions selected to obtain a predetermined natural frequency for the airfoil portion.

Abstract: The invention is intended to reduce the profile loss. For that purpose, according to the invention, a plurality of turbine blades are arranged in the circumferential direction of a turbine driven by a working fluid. Each of the turbine blade is formed such that the curvature of a blade suction surface, which is defined by the reciprocal of the radius of curvature of a blade surface on the blade suction surface side, is decreased monotonously from a blade leading edge defined as the upstream-most point of the blade in the axial direction toward a blade trailing edge defined as the downstream-most point of the blade in the axial direction.

Abstract: Disclosed is a composite heat-dissipating system and its used fan guard which can impart a supercharging function to the heat-dissipating fans of the composite heat-dissipating system for efficient heat dissipation and reduces the noise generated when the heat-dissipating fans are operated. The fan guard includes a frame and a set of guard blades arranged inside and fixed onto an inner surface of the frame. The fan guard can be arranged upstream or downstream of the rotor blades of the heat-dissipating fans and assembled with the heat-dissipating fans in series or in parallel to supercharge the airflow out of the heat-dissipating fans.

Abstract: An annular platform for the nozzle of a low-pressure turbine in a turbomachine, said nozzle having a longitudinal axis and comprising at least one fixed vane disposed downstream from at least one moving blade of a high-pressure turbine, said platform comprising a downstream portion supporting said fixed vane radially defining an aerodynamic channel which extends longitudinally between a leading edge of said fixed vane and a trailing edge of said moving blade, said platform further comprising an upstream portion extending longitudinally beyond the leading edge of said fixed vane towards the trailing edge of said moving blade so as to lengthen said aerodynamic channel.

Abstract: A turbine nozzle includes airfoil and sidewall surfaces. The airfoil and sidewall surfaces have profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Tables I–IV for the pressure and suction sides of the airfoil, and the outer and inner sidewall surfaces, respectively. The X, Y and Z values are distances in inches. The X and Y values for the airfoil, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at the Z distances are joined smoothly with one another to form a complete airfoil shape. The X, Y and Z values of Tables III and IV define the outer and inner sidewall surfaces, respectively, of the gas flowpath.

Abstract: First stage nozzle airfoils have internal core profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein X, Y and Z values are in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define internal core profile sections at each radial distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete internal core profile. The X, Y and Z distances may be scalable as a function of the same constant or number to provide a scaled-up or scaled-down internal core profile. The nominal internal core profile given by the X, Y and Z distances lies within an envelope of ±0.030 inches in directions normal to any internal core surface location.

Abstract: An attachment system for use with a variable incidence vane is disclosed. The attachment system includes a vane arm for joining a unison ring to a vane spindle. The vane arm has an arm portion and a bushing connected to the arm portion. The attachment system further has a pin for joining the vane arm to the unison ring. The pin fits within an interior bore in the bushing and is joined to the unison ring by a dual swage.

Abstract: A turbine blade or vane includes an inner contour profile which is matched to the contour profile of the outer transition region, so that a substantially uniform blade or vane wall thickness is achieved.

Abstract: A guide vane 1, such as may be used in a guide vane stage of a compressor (not shown), the guide vane comprising a main body 2 and a mounting base 4. In one aspect of the invention, at least part 22 of the mounting base 4 is aerofoil shaped. In another aspect of the invention, the mounting base 4 projects outwardly beyond a surface 14 of the main body 2 on one side only of the guide vane 1.

Abstract: A lightweight high temperature bladed turbine disk intended for use in gas turbine engines. The bladed disk comprises a cast integral ring of single crystal airfoils with the primary and secondary crystallographic orientation being the same for each airfoil. Low-angle mismatch boundaries are present in the endwalls that couple adjacent airfoils. The cast ring of single crystal turbine blades is diffusion bonded to a high strength equiaxed disk. The resulting single crystal bladed disk is endowed superior performance, temperature capability, and lower weight and cost, relative to conventional turbines composed of individually cast single crystal blades, which are mechanically inserted into machined slots in the disk, or lower strength cast equiaxed blade rings that are diffusion bonded to a high strength turbine disk.

Abstract: A turbine nozzle segment includes a single hollow airfoil extending radially between radially outer and inner band segments. The airfoil has an airfoil wall with pressure and suction sides extending axially between leading and trailing edges of the airfoil. The airfoil wall surrounds a bifurcated cavity and a bifurcating rib extends through the bifurcated cavity and between the pressure and suction sides of the airfoil wall dividing the bifurcated cavity into forward and aft cavities. A stiffening rib extends radially outwardly from and along a radially outer surface of the outer band segment and axially and circumferentially from a pressure side forward corner of the outer band segment to the bifurcating rib. The outer and inner band segments, the bifurcating rib, and the stiffening rib are integral and made from a unitary one-piece casting.

Abstract: A flow directing device of a gas turbine engine, comprising: an airfoil having a leading edge, trailing edge, suction side and pressure side; a wall abutting the airfoil; and a fillet between the airfoil and wall. The fillet has an enlarged section at the leading edge, along the suction and pressure sides, and towards the trailing edge. The device could be part of a vane segment. In addition to eliminating a horseshoe vortex, the device also reduces heat load on the airfoil by directing the cooler gas from the proximal end of the airfoil to the hotter gas at the medial section of the airfoil.

Abstract: A wind turbine device of increased efficiency is comprised of a set of fixed airfoils that direct wind into a rotor having a plurality of blades. The fixed airfoils may extend to the ground to increase the amount of wind directed into the rotor and may be manufactured from concrete. The rotor blades have a vented portion near the axis of rotation that has been found to increase efficiency for certain blade geometries. For other blade geometries, increased efficiency is observed with no gap at the axis of rotation. The rotor may also be manufactured from composite materials to increase strength while decreasing the moment of inertia for the rotor.

Abstract: A heat-dissipating fan includes a casing having an air outlet, a base mounted in the air outlet, and a plurality of guiding plates in the air outlet. An impeller is mounted on the base that is supported in the air outlet by the guiding plates. The guiding plates are parallel to one another and extend in a direction having an inclining angle with respect to an axial direction of the air outlet. The guiding plates guide airflow and increase wind pressure when the impeller turns.

Abstract: A turbine nozzle includes hollow vanes extending between outer and inner bands. The vanes include a pattern of outlet holes distributed over the pressure and suction sides thereof for discharging cooling air collectively at a reference flowrate. The pattern of holes includes multiple rows of showerhead holes bridging the leading edge, and two rows of gill holes spaced aft therefrom along the suction side. A row of auxiliary holes is spaced aft from the showerhead holes through the pressure side, and the gill holes are sized to counterbalance the added discharge air through the auxiliary holes for maintaining the reference flowrate.

Abstract: A turbine nozzle includes a hollow vane having opposite pressure and suction sides extending in span between outer and inner bands. The vane includes a forward flow channel behind the leading edge, an aft flow channel in front of the trailing edge, and a middle flow channel disposed therebetween. The three flow channels are disposed in flow communication with an outer plenum outside the outer band for receiving cooling air therefrom. The forward and middle channels are also disposed in flow communication with an inner plenum below the inner band for discharging the air. The aft channel discharges air through the inner band outside the inner plenum in split flow at different pressures.

Abstract: A high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the gas flow path of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector the is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, such as on a turbine nozzle. The surface reflects radiation back into the hot gas flow path.

Abstract: A turbine blade/vane includes a profiled blade/vane aerofoil, which extends along a blade/vane center line. A platform is formed on an end region of the blade/vane aerofoil, extending transverse to the blade/vane center line. This is to be designed for a particularly high capability for carrying thermal and mechanical loading. In addition, the arrangement permits reliable cooling with a comparatively small coolant requirement. For this purpose, the platform includes an outer rim which is thickened in comparison with the platform floor. The side wall of the outer rim facing toward the blade/vane center line is slanted relative to the latter. A casting system suitable for manufacturing the turbine blade/vane includes a first shell element which can be positioned in a casting mold, an essentially flat configuration second shell element being displaceable, while being guided in a direction tilted by an angle of more than 10° and of less than 80°, in the first shell element.

Abstract: A method facilitates assembling a turbine nozzle for a gas turbine engine. The method includes providing a turbine nozzle including a plurality of airfoil vanes that extend between an inner band and an outer band, and forming a compound radii fillet that extends between a first of the airfoil vanes and the outer band, such that at least a second of the airfoil vanes is coupled to the outer band only by a single radii fillet.

Abstract: The invention relates to an axial compressor stator for a gas turbine, the stator having a rigid, external, annular frame and axially juxtaposed rings configured within the frame and bearing annuli of stationary vanes. The rings are defined by arcuate segments affixed to the frame. The inside walls of the arcuate segments externally define the aerodynamic conduit for the compressed gaseous fluid. The arcuate segments are brazed segments that include a honeycomb component sandwiched between an inner sheetmetal bounding the aerodynamic conduit and an outer sheetmetal. The outer sheetmetal solely connects the arcuate segments to the frame.

Abstract: Aspects of the invention relate to a method for clocking a row of stationary airfoils relative to an adjacent row of airfoils, and a turbine blade ring configured for use in such a method. The method according to aspects of the invention is advantageous because it can avoid extensive disassembly of the turbine in order to clock the blade ring. The method includes engaging the outer peripheral surface of a blade ring with a plurality of roller pins such that the weight of the blade ring is entirely supported by the roller pins. A tangential force is applied to the blade ring such that the blade ring rotates on the roller pins, thereby adjusting the position of the stationary airfoils affixed to the blade ring. Additional rotational adjustments can be made to the blade ring until the blade ring is clocked to the desired position.

Abstract: An internal coating on an internal passage wall exposed at a passage opening through an article external surface is protected from removal during repair of the article, including removal of at least a portion of an external coating, by a masking assembly disposed about the passage opening. The masking assembly comprises a masking member and a substantially flexible seal, substantially inert to a coating removal medium for the external coating. The masking member is shaped for disposition about the passage opening across a gap between the external surface and the masking member. The substantially flexible seal is disposed across the gap substantially to seal the gap.

Abstract: The present invention relates to an improved system for installing a vane in a gas turbine. The system includes a receptor pocket positioned within a first support structure and a boot attached to a first or inner end of a vane for insertion into the receptor pocket. In a first embodiment of the present invention, the receptor pocket is formed by machining it into the support structure. In a second embodiment of the present invention, the receptor pocket is formed by an insert which is installed into the support structure. In accordance with the present invention, the inner end of the vane having the boot is first inserted into the receptor pocket. After insertion has been completed, the vane is rotated until an outer base is brought into position against an outer support structure. The outer end of the vane is then secured to the outer support structure using studs and nuts.

Abstract: A vane cluster for a turbine engine compressor or turbine includes a shroud 12 with a nonlinear slot 26 extending therethrough to divide the shroud into thermally independent shroud segments 24. The slot is bordered by matching nonlinear surfaces 28 that are easy and inexpensive to produce with conventional wire EDM equipment. The nonlinear profile of the slots effectively resists fluid leakage.

Abstract: An embodiment of a disclosed apparatus for sizing an engine stator has a plurality of radially moveable shoes for engaging the inner and outer rings of the stator. The shoes apply a radially directed force to the inner and outer rings to cause the rings to expand. In particular embodiments, one or more heating mechanisms, such as a ceramic heating blanket, may be used to heat the stator prior to sizing. A method for sizing an engine stator also is disclosed. The method comprises simultaneously displacing the inner and outer rings of the stator generally radially outwardly so as to increase the respective diameters of the inner and outer rings.

Abstract: A turbine nozzle includes a plurality of vanes joined at opposite ends to outer and inner bands. The inner band has a forward hook which is segmented to reduce thermal mismatch. And, in additional embodiments the vane includes an impingement baffle having preferential cooling.