Abstract: A gas turbine airfoil (10) includes a serpentine cooling path (32) with a plurality of channels (34,42,44) fluidly interconnected by a plurality of turns (38,40) for cooling the airfoil wall material. A splitter component (50) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel (46) passing in between the outer wall (28) and the inner wall (30) of the pressure side (24) and a suction-side channel (48) passing in between the outer wall (28) and the inner wall (30) of the suction side (26) longitudinally downstream of an intermediate height (52). The cross-sectional area of the pressure-side channel (46) and suction-side channel (48) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.

Abstract: A turbine blade including an airfoil, the airfoil including an airfoil outer wall having pressure and suction sidewalls joined together at chordally spaced apart leading and trailing edges extending radially outwardly from a blade root to a blade tip surface. A continuous squealer tip rail extends radially outwardly from and substantially continuously around the blade tip surface forming a radially outwardly open squealer pocket. The squealer tip rail includes an aft portion adjacent to the trailing edge, where the pressure side tip rail is discontinued in the chordal location of the aft portion. A plurality of chordally spaced film cooling holes are provided in a surface of the suction sidewall and provides a cooling fluid flow along the aft portion of the squealer tip rail surface to define a sub-boundary layer or film of cooling fluid between a hot gas vortex flow and the aft portion of the squealer tip rail at the suction sidewall.

Abstract: A hybrid ceramic matrix composite (CMC) structure 10 and method for fabricating such an structure are provided. A CMC substrate 12 includes layers 16, 18, 20 of ceramic fibers. Fugitive objects 22 are disposed on at least one of the plurality of layers prior to laying a subsequent layer of ceramic fibers. An outer surface of the subsequent layer influences a shape of the outer surface of the substrate by defining protuberances 24 on the outer surface of the substrate where respective cavities 26 are formed beneath respective protuberances upon dissipation of the fugitives. A liquefied ceramic coating 34 is deposited on the outer surface of the ceramic substrate to fill the cavities. When the ceramic coating is cured to a solidified state, the cavities containing the solidified coating constitute an anchoring arrangement between the ceramic substrate and the ceramic coating.

Abstract: A turbine component tracking system is provided. The turbine component tracking system is advantageously adapted to determine the remaining life of individual turbine components based on how and where they are used as well as if and how they are repaired. The turbine system is also advantageously adapted to track and analyze design, manufacturing and repair changes or modifications performed on turbine components.

Abstract: A process for producing a stable high-temperature catalyst for reduction of nitrogen oxides in combustion exhaust gases at operating temperatures from 300° C. to over 700° C. without the need for exhaust dilution. A zeolite material is steam-treated at a temperature and duration sufficient to partially de-aluminize the zeolite to approximately a steady state, but not sufficient to fully collapse its chemical structure. Iron is added to the zeolite material. The zeolite material is calcined at a temperature, humidity, and duration sufficient to stabilize the zeolite material. Examples and specifications for ranges, order, and durations of steaming, calcining, and other steps are provided.

Abstract: A gas turbine engine fuel nozzle assemblage (10) comprises a nozzle outer body (25) comprising a premixing section (23) and an interiorly disposed first fuel passageway (22) to supply fuel to the premixing section (23), and more interiorly a central channel (33) comprising an inlet (30) from an air supply, an aperture (19) at an upstream end to receive a lance (18), and a downstream-disposed annular bearing surface (21) adapted to slidingly engage the lance, the fuel nozzle assemblage adapted so that the lance may be alternatably selected as any interchangeable one of a gas lance, an oil lance, and a non-fueling dummy lance. Between the lance and the central channel (33) is formed a cooling fluid passageway (24) that is in fluid communication with the inlet (30) and one or more cooling fluid passageway exits (60). The interchangeability of the lances (18) is effective to provide a desired flexible functionality to said fuel nozzle assemblage.

Abstract: A method of remotely monitoring the radiant energy (6) emitted from a turbine component such as a turbine blade (1) having a low-reflective surface coating (3) which may be undergoing potential degradation is used to determine whether erosion, spallation, delamination, or the like, of the coating (3) is occurring.

Abstract: A combustion catalyst coating (36) applied to the surface of a ceramic thermal barrier coating (34) which is supported by a metal substrate (32). The microstructure of the thermal barrier coating surface provides the necessary turbulent flow and surface area for interaction of the catalyst and a fuel-air mixture in a catalytic combustor of a gas turbine engine. The temperature gradient developed across the thermal barrier coating protects the underlying metal substrate from a high temperature combustion process occurring at the catalyst surface. The thermal barrier coating deposition process may be controlled to form a microstructure having at least one feature suitable to interdict a flow of fuel-air mixture and cause the flow to become more turbulent than if such feature did not exist.

Abstract: The present invention provides for a static excitation system for a superconducting rotor that comprises multiple brushes (12) in contact with the superconducting rotor winding (6), normally via collector rings and field leads. A power conditioning device (16) is connected to the brushes (12), and an energy storage device (18) is linked to the power conditioning device (16). The power conditioning device provides power from the energy storage device to the superconducting rotor when required, and when power to the superconducting rotor is not required, the power conditioning device takes excess power from the superconducting rotor and stores it in the energy storage device.

Abstract: A catalytic combustor (14) includes a first catalytic stage (30), a second catalytic stage (40), and an oxidation completion stage (49). The first catalytic stage receives an oxidizer (e.g., 20) and a fuel (26) and discharges a partially oxidized fuel/oxidizer mixture (36). The second catalytic stage receives the partially oxidized fuel/oxidizer mixture and further oxidizes the mixture. The second catalytic stage may include a passageway (47) for conducting a bypass portion (46) of the mixture past a catalyst (e.g., 41) disposed therein. The second catalytic stage may have an outlet temperature elevated sufficiently to complete oxidation of the mixture without using a separate ignition source. The oxidation completion stage is disposed downstream of the second catalytic stage and may recombine the bypass portion with a catalyst exposed portion (48) of the mixture and complete oxidation of the mixture.

Abstract: A printed circuit board (PCB 22) capable of withstanding ultra high G forces and ultra high temperature as in a gas turbine (11). The PCB includes a substrate having a plurality of cavities (30A, 36A) formed therein for receiving components of a circuit, and conductors embedded in the PCB for electrically connecting the components together to complete the circuit. Each of the cavities has a wall (36A?) upstream of the G-forces which supports the respective component in direct contact in order to prevent the development of tensile loads in a bonding layer (37A). When the component is an integrated circuit (50), titanium conductors (63) are coupled between exposed ends of the embedded conductors and contact pads on the integrated circuit. A gold paste (51) may be inserted into interstitial gaps between the integrated circuit and the upstream wall.

Abstract: A fuel injector (36) for alternate fuels (26A, 26B) with energy densities that differ by at least about a factor of two. Vanes (47B) extend radially from a fuel delivery tube structure (20B) with first and second fuel supply channels (19A, 19B). Each vane has first and second radial passages (21A, 21B) communicating with the respective fuel supply channels, and first and second sets of apertures (23A, 23B) between the respective radial passages and the surface (49) of the vane. The first fuel supply channel, first radial passage, and first apertures form a first fuel delivery pathway providing a first fuel flow rate at a given backpressure. The second fuel supply channel, second radial passage, and second apertures form a second fuel delivery pathway providing a second fuel flow rate that may be at least about twice first fuel flow rate at the given backpressure.

Abstract: An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.

Abstract: A ceramic ring segment for a turbine engine that may be used as a replacement for one or more metal components. The ceramic ring segment may be formed from a plurality of ceramic plates, such as ceramic matrix composite plates, that are joined together using a strengthening mechanism to reinforce the ceramic plates while permitting the resulting ceramic article to be used as a replacement for components for turbine systems that are typically metal, thereby taking advantage of the properties provided by ceramic materials. The strengthening mechanism may include a ceramic matrix composite overwrap or plurality of overwraps designed to help prevent delamination of the ceramic plates when the ceramic article is in use by placing the plates in compression.

Abstract: An inner mounting ring (20) for gas turbine flow path components such as shroud ring segments (24). The inner ring (20) may be mounted to an outer ring (22) on radially slidable mounts (26, 28) that maintain the two rings (20, 22) in coaxial relationship, but allows them to thermally expand at different rates. This allows matching of the radial expansion rate of the inner ring (20) to that of the turbine blade tips (32), thus providing reduced clearance (33) between the turbine blade tips (32) and the inner surface of the shroud ring segments (24) under all engine operating conditions. The inner ring (20) may be made of a material with a lower coefficient of thermal expansion than that of the outer ring (22).

Abstract: An inspection apparatus (10) applying two dimensional nondestructive examination images onto a three dimensional solid model of a component (12) to display a virtual component (73) that may be manipulated to perform a nondestructive inspection. The two dimensional nondestructive examination images may be acquired from a plurality of views of the component in order to provide full coverage of the surface to be inspected, with appropriate stitching of images in regions of overlap between adjacent views. The two dimensional images (62) may be color or black and white photographs or ultraviolet or infrared images, for example. Multiple types of nondestructive examination images, results of inspection data evaluations, and design, operational and/or maintenance information may be displayed separately or jointly on the three dimensional solid model.

Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: A blade gap control system configured to move a blade ring of a turbine engine relative to a blade assembly to reduce the gaps between the tips of the blades and the blade rings to increase the efficiency of the turbine engine is provided. The blade rings can be at an acute angle with respect to the rotational axis of the blade assembly. The axial movement of the blade ring can be done by a pressure differential supplied across the blade ring, the thermal expansion and/or contraction of a linkage or by a piston.

Abstract: A method of evaluating a power plant having a design life based on operating the plant within an allowable chemical exposure range includes accumulating a history of a chemical exposure of a steam generating portion of the power plant. The method also includes determining a remaining life of the plant based on the history of the chemical exposure and assuming continued operation of the plant within the allowable chemical exposure range. The method may also include evaluating an economic value of operating the plant based on the remaining life of the plant.

Abstract: A seal assembly limits gas leakage from a hot gas path to one or more disc cavities in a gas turbine engine. The seal assembly includes a seal apparatus associated with a blade structure including a row of airfoils. The seal apparatus includes an annular inner shroud associated with adjacent stationary components, a wing member, and a first wing flange. The wing member extends axially from the blade structure toward the annular inner shroud. The first wing flange extends radially outwardly from the wing member toward the annular inner shroud. A plurality of regions including one or more recirculation zones are defined between the blade structure and the annular inner shroud that recirculate working gas therein back toward the hot gas path.