Abstract: A gas turbine airfoil (20) having a load-bearing core (30). A honeycomb structure (40A, 42A) is attached to pressure and/or suction sides (22, 24) of the core and is filled with ceramic insulation (50). A ceramic matrix composite boot (60A, 60B, 60C) may cover the leading edge (26) of the core. Edges (61, 62) of the boot may be attached to the core by rows of pins (63A, 63B) or by flanges (65) inserted in slots (69) in the core. The pins may be formed in place by forming pin holes (64) in the boot, clamping the boot onto the core, filling the pin holes with metal or ceramic and metal particles, and heating the particles for internal cohesion and solid-state diffusion bonding (66) with the core. The boot may have a central portion (71) that is not bonded to the core to allow differential thermal expansion.

Abstract: A gas turbine case is provided including an outer case surface, and a channel portion formed as a recessed area extending radially inwardly into the outer case surface. The channel portion extends about a circumference of the case. An outer flow jacket is attached to the outer case surface and extends over the channel portion to define an enclosed cooling passage along the outer case surface. At least one inlet passage and at least one outlet passage are provided in fluid communication with the enclosed cooling passage to convey air to and from the cooling passage.

Abstract: A process for producing a component of a gas turbine having a substrate with a metallic layer is provided. The metallic layer is a MCrAlX layer which is treated at temperatures elevated above the operating temperature, by at least 50° C., so that the oxidation and corrosion behavior are improved. In particular a MCrAlX of the type, NiCoCrAlX is used.

Abstract: An apparatus, especially a steam turbine, including a first delimiting part and a second delimiting part is provided. The delimiting parts are attached to one another forming a joint and enclosing at least a part of a first pressure chamber. A shielding element is also provided on the sides of the delimiting parts facing towards the first pressure chamber and is arranged to form a seal completely covering the joint, so that a cavity is formed between the delimiting parts and the shielding element. A line is routed in the cavity connecting the cavity to a second pressure chamber. In addition, a method for decreasing the force acting on a joint, which is formed by the joining together of a first delimiting part and a second delimiting part of an apparatus, especially a steam turbine, and for reducing the attachment forces acting on the joint is provided.

Abstract: A turbine (100) of a turbine installation, especially a steam turbine of a steam turbine installation, includes at least one radial or diagonal turbine stage (120) with radial or diagonal inflow and axial outflow, and also at least one axial turbine stage (121-125) with axial inflow and axial outflow. The at least one radial or diagonal turbine stage (120) forms the first stage of the turbine (100) and the at least one axial turbine stage (121-125) is arranged downstream of the radial or diagonal turbine stage (121) as an additional stage of the turbine. The at least one radial or diagonal turbine stage (120) has a higher temperature resistance than the at least one axial turbine stage (121-125). The turbine (100) makes it possible to significantly increase the process temperature of the steam turbine installation, wherein measures for increasing the temperature resistance need only to be adopted for components of the radial or diagonal turbine stage (120).

Abstract: A tungsten bronze structured ceramic material as a thermal barrier coating is described wherein the tungsten bronze structured ceramic coating material has the formula AO—BvOw—CyOz where O stands for Oxygen, A stands for a 2+ or a 1+ cation, B stands for a 2+ or 3+ cation and C stands for a 4+ or a 5+ cation. The thermal barrier coating may be applied for gas turbine components.

Abstract: A turbocharger is provided having a bearing housing (100) and a heat shield (200) staked thereto. The staking process can be a compression staking process that deforms one or more portions of the bearing housing (100). The bearing housing (100) can have an annular channel (150) and the compression staking can be done to an outer edge (160) of the annular channel (150) to form one or more compression projections (400). The size and shape of the annular channel (150) can be substantially the same as the size and shape of an outer portion of the heat shield (200) to allow for a press-fit therebetween.

Abstract: A shroud suitable for use in a gas turbine engine exhibits substantially uniform thermal growth.

Abstract: A film cooling structure 10 includes a structural wall 11 that has an outer surface 12 exposed to combustion gas and an inner surface 13 positioned opposite to the outer surface 12, and film cooling holes 14 are formed at the structural wall 11 and introduce cooling air from the inner surface 13 toward the outer surface 12 in order to perform film cooling. The film cooling hole 14 includes an introducing portion 14a that extends to a middle position in the structural wall 11 from the inner surface 13 toward the outer surface 12, an enlarged portion 14b of which the cross-sectional area is gradually increased toward the outer surface 12 from an end of an outer surface side of the introducing portion 14a and which is opened at the outer surface 12, and a partition portion 16 that partitions the inside of the enlarged portion 14b into a plurality of spaces in a width direction of the hole perpendicular to a flow direction of the combustion gas.

Abstract: A steam turbine 10 is comprised of a double-structured casing configured of an outer casing 21 and an inner casing 20, a turbine rotor 23 disposed through the inner casing and having a plurality of stages of moving blades 22 implanted, and a plurality of stages of stationary blades 25 disposed alternately with the moving blades 22 in the axial direction of the turbine rotor 23 in the inner casing 20. The steam turbine 10 is further provided with a discharge passage 30 which externally guides steam, which has flown in the inner casing and passed the final stage moving blades while performing expansion work, directly from the inner casing interior.

Abstract: A process for the formation of positive features on the surface of a turbine shroud component is described. The process involves applying a feature-forming material to a selected portion of the component surface with a laser consolidation apparatus, according to a pre-selected shape and size for the positive features. A gas turbine engine, comprising a shroud component which contains positive features formed according to embodiments of this process, represents another embodiment of this invention. Methods for modifying the shape of at least one positive feature on a surface of a shroud component are also described.

Abstract: A system for ventilating a downstream cavity of an impeller of a centrifugal compressor in a turbomachine, this cavity being delimited by the downstream face of the impeller and by an annular end-piece of a diffuser and being ventilated by taking air from the outlet of the compressor, the system comprising at least one annular deflector mounted in this cavity in order to oppose the rise of hot air along the downstream face of the impeller and in order to cause the diverted hot air to be driven by the air taken from the outlet of the compressor.

Abstract: The present invention provides a coated article made using a plating process, either electroless or electrolytic, whereby nickel is plated on the article using a solution that includes a suspension of powders or containing one or more of the following elements: Ni, Cr, Al, Zr, Hf, Ti, Ta, Si, Ca, Fe, Y and Ga. Optionally, the coating is then heat treated at a temperature above about 1600.degree. F. for an effective amount of time to allow to homogenize the coating by allowing effective interdiffusion between the species. The level of aluminum may be altered to produce a coating of the predominantly beta phase of the NiAl alloy composition. Optionally, a TBC layer is applied over the predominantly beta phase NiAl alloy composition metallic bond coat.

Abstract: A turbine vane of a gas turbine, especially a gas turbine aircraft engine, is disclosed. The turbine vane includes a vane base body with an outer surface forming a suction side and a pressure side, the outer surface of the vane base body being at least partially coated with a thermal barrier coating. The thermal barrier coating extends continuously or uninterruptedly at least largely over the suction side and largely over the pressure side of the surface of the vane base body, with the layer thickness of the thermal barrier coating being variable or adjustable.

Abstract: According to an embodiment of the invention, an article of manufacture for use in a gas turbine engine is disclosed. The article comprises a part having a surface covered with a ceramic thermal barrier coating. The thermal barrier coating has an outer surface covered with a sacrificial phosphate coating, wherein the sacrificial phosphate coating reacts with contaminant compositions to prevent contaminant infiltration into the thermal barrier coating.

Abstract: A fluid cooled fastener assembly for use in a high temperature environment and a method of fluid cooling a fastener assembly are provided. The fastener has a coolant passage extending axially through the fastener and a coolant collector coupled to an end of the fastener. The coolant collector includes a collector opening configured to capture a portion of a cooling fluid flowing past the coolant collector and a contoured passage in communication with the collector opening and the passage in the fastener. The contoured passage is configured to accelerate the cooling fluid captured by the coolant collector and to direct it into the coolant passage in the fastener.

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

Abstract: A gas turbine engine includes a housing with a duct wall that defines a generally annular and axially elongated hot gas flow path for passage of combustion gas. The engine further includes a contoured shroud mounted within the internal engine cavity and defining a hot gas recirculation pocket for receiving hot gas ingested from the hot gas flow path through the annular space and for recirculating the ingested hot gas back through the annular space to the hot gas flow path. The contoured shroud includes a base wall extending radially inwardly from the duct wall, an inboard wall extending from the base wall in an axial direction toward the rotor, and an end wall extending from the inboard wall in a radially outward direction. The end wall terminates in a circumferentially extending free edge disposed in proximity to the annular space. The end wall defines an opening.

Abstract: One embodiment of a heat exchange system is disclosed having a heat exchange compressor and multiple evaporators capable of operating at different heat transfer requirements. The heat exchange compressor may be a single-stage or multi-stage compressor. In one form the evaporators return working fluid in separate streams to the heat exchange compressor. The heat exchange compressor may be integrated with a gas turbine engine and includes a number of inlets that correspond to a number of separate evaporators. Each inlet can be configured to receive working fluid at different locations within a pressure and velocity flow field created in the compressor. The heat exchange compressor may be driven by a shaft of the gas turbine engine and may be positioned at a variety of locations.

Abstract: An outlet guide vane assembly for turbomachines is provided. The outlet guide vane assembly comprises one or more outlet guide vanes, wherein each of the one or more outlet guide vanes comprises a first surface and a second surface, and wherein the one or more outlet guide vanes are disposed between an inner wall and an outer wall of an engine and a surface cooler layer disposed on at least a portion of the first surface, the second surface, or both of the one or more outlet guide vanes, wherein the surface cooler layer comprises a metal foam, a carbon foam, or a combination thereof, wherein the metal foam, the carbon foam or the combination thereof is configured to augment heat transfer and enhance acoustic absorption.

Abstract: The present invention is a hover infrared suppression system for a gas turbine engine comprising a hover infrared suppression system having an upstream first stage, a second stage downstream of the first stage and a third stage downstream of the second stage, the engine operating at a temperature sufficient to cause the hover infrared suppression system to emit infrared radiation. The present invention further comprises a high reflectivity coating applied over a preselected area of at least one of the stages of the hover infrared suppression system to reduce the infrared radiation emitted from the engine, the high reflectivity coating being fired after application.

Abstract: A turbine nozzle includes a hollow vane mounted between inner and outer bands. The inner band includes a mounting flange between forward and aft lips. An aft pocket is found in the inner band between the flange and aft lip. And, an impingement bore extends through the flange into the windward half of the pocket and is directed aft toward the opposite leeward half of the pocket for co-rotation with purge flow during operation.

Abstract: A gas turbine engine includes an engine core extending along a core axis, a cooling-air delivery duct on the engine core, and a removable nacelle cowl overlying the engine core. A cooling-air nacelle-cowl duct delivers cooling air to the cooling-air delivery duct. At least a portion of the length of the cooling-air nacelle-cowl duct is integral with the nacelle cowl and not directly supported on the engine core.

Abstract: A method of cooling a double flow steam turbine includes supplying steam flow to each nozzle of the sections of the turbine; reversing a portion of each steam flow to provide a reverse steam flow from an aft side to a forward side of each section. Each reverse steam flow is directed to an annular space between a rotor and a tub. The method further includes removing the reverse steam flows through a pipe, the pipe having a first end at the annular space at a first pressure and a second end at a second pressure that is lower than the first pressure. A double flow steam turbine, includes a pair of nozzles, each nozzle being provided at a section of the turbine; a rotor supporting buckets of the sections; a tub supporting the pair of nozzles; and a pipe extending from an annular space between the tub and the rotor. The pipe has a first end at the annular space and second end. A pressure at the first end of the pipe is greater than a pressure at the second end.

Abstract: Systems and methods for cooling heated components in a turbine are provided. According to one embodiment, a system for cooling a turbine is provided that may include at least one liquid source which may include a coolant liquid. The system may also include at least one liquid nozzle in fluid communication with the liquid source or sources and operable to deliver the coolant liquid in an atomized form adjacent to at least one heated turbine component positioned in a hot gas path of the turbine. Upon delivering the atomized coolant liquid adjacent to the heated turbine component or components, at least a portion of the coolant liquid substantially changes phase to a gas.

Abstract: The invention relates to the use of a thermal insulating layer for a housing of a steam turbine in order to even out the deformation behaviour of different components based on different heatings of the components.

Abstract: A turbine shaft may be used to remove a heat load from pliant bearings without requiring the use of a process fluid for cooling the turbine shaft. The turbine shaft comprises a heat conductive sleeve disposed between an outer surface of a tie rod shaft and an inner surface of a bearing journal; the heat conductive sleeve having a sleeve inner surface separated from a sleeve outer surface by a sleeve thickness; the heat conductive sleeve having a first end separated from a second end longitudinally about a center axis; the sleeve outer surface being in physical contact with the inner surface of the bearing journal; and the heat conductive sleeve having a thermal conductivity that is greater than a thermal conductivity of the bearing journal.

Abstract: A gas turbine (1) includes a rotor (2) which has two rotor blade rows (5) with a plurality of rotor blades (6), and also a rotor heat shield (7), which is arranged between them, with a plurality of heat shield elements (12), and with a stator (3) which has a stator blade row (8), with a plurality of stator blades (9), which is arranged axially between the two adjacent rotor blade rows (5). The stator blades (9) have a stator sealing structure (10) radially on the inside. The heat shield elements (12) have a rotor sealing structure (13) radially on the outside which interacts with the stator sealing structure (10) for forming an axial seal (14). Furthermore, a blade radial seal (15) is formed between two adjacent rotor blades (6), and also a heat shield radial seal (16) is formed between two adjacent heat shield elements (12), and in each case separates a gas path (17) from the rotor (2).

Abstract: Turbine components with different types of coatings on different parts thereof are described. The coatings are chosen such that they are especially adapted to the thermal and corrosive conditions being present on the parts of the component during use. A method to coat a turbine component is also described.

Abstract: A turbocharger comprising a cartridge which is arranged in a turbine casing for variable turbine geometry; a bearing housing which is arranged between the turbine casing and a compressor housing of a compressor impeller and in which a bearing assembly is arranged for a shaft which supports the turbine wheel and the compressor impeller; and a spring washer which is arranged between the cartridge and the bearing housing, the spring washer being composed of at least two material layers.

Abstract: The present invention relates to an exhaust gas turbocharger for a motor vehicle, comprising a turbine housing having a first area in which exhaust is carried to a turbine wheel and in which the turbine wheel is essentially situated and having a second area connected axially to the first area, such that the first area and the second area are essentially separated from one another by a separating element. It is provided that for thermal insulation, the separating element (14) is made of a material having a low thermal conduction (24) and/or thermal capacity.

Abstract: The invention relates to the use of a thermal insulating layer for a housing of a steam turbine in order to even out the deformation behaviour of different components based on different heatings of the components.

Abstract: A method for assembling a gas or steam turbine is provided. The method includes providing an insulator and positioning the insulator between a vane support and a vane such that the insulator facilitates preventing hot gas migration into the vane, and such that during operation, hot gas is channeled from a high pressure side of the vane to a low pressure side of the vane. A vane assembly for a turbine rotor assembly is also provided. The vane assembly includes a vane support and an insulator including a projecting portion. The assembly also includes a vane. The insulator is coupled to the vane support such that the projecting portion is between the vane and a nozzle support strut to facilitate hot gas flow from a pressure side of the projecting portion to a suction side of the projecting portion.

Abstract: The invention relates to a blower apparatus with an electric motor, which drives a blower wheel, is arranged in a motor housing and has a commutator with at least two brushes, which form in each case an associated brush bearing face with respect to the commutator, wherein a cooling air flow for cooling the electric motor is diverted into a cooling air channel of the motor housing by a blower air flow produced by the blower wheel. The invention is based on the object of designing the motor cooling system of the blower apparatus of the generic type in such a way that it is more efficient. The object is achieved by virtue of the fact that the cooling air channel has an air distributor, which splits the cooling air flow into at least two cooling air subflows, which are in each case supplied to one of the brush bearing faces directly and in targeted fashion.

Abstract: A turbocharger for internal combustion engines, having a heat shield disposed between a turbine housing and center housing of the turbocharger for shielding the center housing from exhaust gases passing through the turbine. The heat shield in one embodiment is centered on a generally conical surface of the center housing, by engagement between the conical surface and a centering portion of the heat shield. In another embodiment, the heat shield has a generally conical portion that is non-axisymmetric, having a plurality of radially enlarged lobes that engage a radially inwardly facing surface of the turbine housing for centering the heat shield.

Abstract: The invention relates to a turbomachine module comprising a device to improve radial clearances, comprising an outer casing (22), an inner casing (21) and at least one shock absorbing ring (5) connecting the casings, the module also comprising an annular cavity (43) on the downstream side of the ring (5), with a passage (42) formed in it through which an airflow (4) is drawn off from the primary flow (Fp). According to the invention, the device to improve radial clearances comprises a thermal inertia control device for the shock absorbing ring mounted on the shock absorbing ring (5), downstream from it, this device comprising at least one coat of thermal insulation (8), and at least partly delimiting the annular cavity (43) on the downstream side of it (43).

Abstract: A fan includes a frame, a motor and an impeller. The motor is disposed on the frame and coupled to the impeller to drive the impeller to rotate. The impeller includes a hub, a plurality of blades and an air exhausting structure. The hub has an opening formed on a top portion of the hub. The blades are disposed around the hub. The air exhausting structure is disposed in the opening for discharging hot airs generated by the motor out of the hub.

Abstract: A sealing structure includes a labyrinth seal provided between a housing and an outer periphery of an end portion of a turbine shaft on a turbine wheel side, and a piston ring seal disposed on a side toward a center of the turbine shaft in an axial direction thereof from a position on the turbine shaft at which the labyrinth seal is provided, and a ring groove is formed along a circumferential direction of the housing in a middle portion of the labyrinth seal to reverse a flow direction of exhaust gas entering from the exhaust passage.

Abstract: A feed pump with a variable-speed drive for adjustably metering a feed quantity, which is designed as a single-stage centrifugal pump with a radial wheel arranged to rotate, without a sealing gap, in an impeller chamber of a pump casing, to convey a fluid between a pump inlet and outlet. The radial wheel is connected to a variable speed drive motor having a five-digit rotational speed range, receives the flow centrally, is provided with feed ducts, and has an outside diameter of up to 50 mm, while the centrifugal pump is designed for partial-load operation, the feed quantities of which range from 0 to 3600 ml/min with lifts of 20 to 300 meters. The inside diameter of the impeller chamber is at most 4% larger than the outside diameter of the radial wheel. A seal is arranged between the impeller chamber and the radial wheel or its shaft, and the circumference of the impeller chamber is provided with one or more outlet ducts arranged at an acute angle or tangentially to the outside diameter of the radial-wheel.

Abstract: A turbine casing may include an outer surface with a false flange and an inner surface with a heat sink positioned adjacent to the false flange.

Abstract: The turbine casing as described herein may include a first section flange, a second section flange, the first section flange and the second section flange meeting at a joint, and a heat sink positioned about the joint.

Abstract: A cooling fan (100) includes a frame (10), a circuit board (20) and a rotor (30). The frame includes a housing (12) and a plate (14) connecting with the housing. The circuit board is mounted on the plate. The rotor is received in the housing and includes a hollow hub (32) and a plurality of blades (38) extending from the hub. A plurality of light emitting diodes (40) is mounted on the circuit board. When the rotor rotates, each LED projects light towards the rotor and to the frame so that the light will be reflected and deflected by the fan whereby the whole fan glows and produces a fantastic visual effect.

Abstract: Protection device (10) for a stator of a turbine of the type comprising a series of annular sectors (12) constrained to each other by means of connection means, each sector of said series of sectors (12) comprises a first surface (13), suitable for contacting the stator which has at least one cavity (14) for the cooling of the corresponding sector, and a second surface (17) which faces a rotor of the turbine. In a rest configuration in which the turbine is not operating, the second surface (17) of each sector (12) has an axial section having an eccentricity (40) with respect to the trace of the rotation axis of the turbine, whereas in an operating configuration in which the turbine is operating under regime conditions, the second surface (17) of each sector (12) has an axial section which is centered with the trace of the rotation axis of the turbine.

Abstract: A turbine nozzle segment may have a band having a flange extending radially from a non-flowpath side and an aft end. A plurality of airfoils may extend radially from a flowpath side of the band and may have trailing edges. A plurality of cooling holes may be disposed in the flange and directed at the aft end between the trailing edges.

Abstract: A cooling system for a turbine airfoil of a turbine engine having a trailing edge cooling channel with bifurcated exhaust channels formed by suction and pressure side trailing edge cooling channels in fluid communication with a central trailing edge cooling channel. The suction and pressure side trailing edge cooling channels may be separated with a trailing edge rib. The suction and pressure side trailing edge cooling channels may be recessed from the airfoil external surface to control the flow of cooling fluids from the cooling system such that the exhaust flow minimizes shear mixing and thus lowers the aerodynamic loss yet maintains high film cooling effectiveness for the airfoil trailing edge.

Abstract: Systems, methods, and devices are disclosed, including a heat exchanger having a manifold with a coolant inlet and a coolant outlet and a plurality of tubes with interiors fluidly connected to the manifold. In some embodiments, the plurality of tubes and the manifold are configured to fit within a bonnet of a compressor.

Abstract: Cooling air flow in a gas turbine engine, for example to a first turbine rotary stage is modulated in accordance with a predetermined cooling schedule by means of a fluidic valve arrangement disposed in the interior of the engine. Since fluidic valves have no moving parts they are more reliable than conventional mechanical and electrically operated valves and able to withstand better the environment in the interior of a gas turbine engine. The fluidic valves are operated remotely by a relatively small control flow that may be governed by an electrically operated valve mounted externally and under the control of a predetermined cooling schedule.

Abstract: A turbine vane for a gas turbine engine having an outer wall of nonuniform thickness. The turbine vane may be formed from a generally elongated airfoil formed from an outer wall having a leading edge, a trailing edge, a pressure side, a suction side, a first endwall at a first end, a second endwall at a second end opposite the first end, and an internal cooling system positioned internally of the outer wall. The outer wall may be formed of a non-uniform thickness such that aspects of the outer wall positioned between an outboardmost portion of the outer wall and an inboardmost portion of the outer wall are thinner than the outboardmost and inboardmost portions of the outer wall.

Abstract: A hydrodynamic machine, and especially a hydrodynamic coupling, for a turbine compound system includes (a) a bladed primary wheel and a bladed secondary wheel which together form an operating space which is filled or can be filled with operating medium, and (b) a housing which encloses the primary wheel and the secondary wheel and which is manufactured from sheet metal. The housing is manufactured from at least two separate sheet metal components which are placed on the primary wheel and/or the secondary wheel, are aligned with each other at a joint abutment location and are welded together. In the area of the welded joint abutment, a relief which is open toward the joint abutment is provided and which extends along the joint abutment in the surface of the primary wheel and/or the secondary wheel.

Abstract: A turbojet turbine or compressor stage, comprising at least one rotor disk mounted inside a sectorized ring (120) supported by a casing (116) and comprising a circumferential flange (132) radially clamped on an annular rail (136) of the casing by a substantially C-shaped cross-section annular locking device (138) which is axially engaged on the casing rail and on the flange of the ring, a substantially cylindrical heat shielding foil (146) being interposed between the ring and the device, the foil, or the device, comprising at least one extra thickness (150) which respectively bears on the device, or the foil, to radially immobilize this foil.