Abstract: A cooling hole having an inlet passage forming an inward spiral flow path and an outlet passage forming an outward spiral flow path in which the two paths are counter flowing in order to improve the heat transfer coefficient. The spiral cooling hole is used in a blade outer air seal (BOAS) for a turbine in which the edges of the shroud segments include a counter flowing micro serpentine flow cooling circuit with thin diffusion discharge cooling slots for the BOAS edges. The total BOAS cooling air is impingement from the BOAS cooling air manifold and metered through the impingement cooling holes to produce impingement cooling onto the backside of the BOAS. The spent cooling air is then channels into the multiple micro serpentine cooling flow circuits located around the four edges of the shroud segments.

Abstract: A turbine blade with a squealer pocket formed by a tip rail extending along the pressure side and suction side of the tip. A row of diffusion notches are spaced along the inner sides of the pressure side tip rail and the suction side tip rail, each notch formed by a peak and a valley and opening into the pocket to function as a diffuser. Each notch is supplied with cooling air through a tip convective cooling hole that opens into the bottom of each notch. The pocket floor is without tip cooling holes so that the cooling air discharged into the notches function to push away the vortex flow that would form along the inner side of the tip rails to improve the cooling effectiveness and reduce the tip rail metal temperature.

Abstract: The present invention is a process for designing the internal cooling passages with these small scale features using a meshed solid model in a FEA program to perform the CFD analysis. Instead of modeling these small scale features (cooling holes, impingement holes, and turbulators) in the meshed FEA solid model, most of the small scale features are replaced by grid extraneous source terms in which the small scale feature is eliminated from the solid model and replaced with point source terms. The source terms can duplicate the effect of the feature within the analysis without requiring the complex analysis that such features would require in the analysis. Text files for each of the cooling holes, impingement holes and turbulators are inputted and then translated into grid extraneous source terms that include position, energy, and continuity. A conjugate CFD solver performs an analysis and produces new values for the thermal and boundary conditions.

Abstract: A diffuser for a turbo machine, such as an exhaust diffuser for an industrial gas turbine engine, the diffuser including a plurality of struts connecting the inner shroud of the diffuser to the outer shroud, where the struts have trailing edges with a step shape to reduce the losses and improve the performance of the diffuser. The stepped shaped trailing edge includes three or more steps with the middle step being the longest in the chordwise direction and the remaining steps being symmetric about the chord of the strut.

Abstract: A film cooling hole for a turbine airfoil used in a gas turbine engine, where the film cooling hole includes an expansion section with a downstream side wall having a first expansion in the middle section and a second expansion greater than the first expansion in the two corners of the downstream side wall. The corners with a greater expansion seal off hot gas ingestion between the film flow and the airfoil surface immediately downstream from the hole opening which minimize the vortices' formation and thus improve the film effectiveness level. The film cooling hole includes two side walls with an expansion of around 10 degrees and an upstream side wall with no expansion. An inlet to the diffusion section includes a metering section of constant diameter to meter cooling air flow.

Abstract: A turbine rotor blade formed from a main support spar and having a thin thermal skin bonded or formed to the support spar to form the outer airfoil surface of the blade. The main support spar is formed from a plurality of ribs arranged to form a cooling air supply cavity near the leading edge region, a plurality of impingement chambers and a plurality of spent air collector chambers arranged along the pressure side and suction side of the airfoil to form a series of impingement chambers and collector chambers to provide near wall cooling for the thermal skin. A row of exit cooling holes located along the trailing edge discharges the cooling air from the last collector chamber. In another embodiment, some of the collector chambers are connected to rows of film cooling air to discharge film cooling onto the outer airfoil surface. The thermal skin includes micro pin fins on the inner wall to enhance the cooling effect of the impinging cooling air.

Abstract: A hydrocarbon presence sensor used to detect for the presence of a fuel in a buffer cavity of an inter-propellant seal. A sensor is formed on a rotor shaft in the buffer cavity and an optical imaging device is used to view a change in appearance of the sensor when a hydrocarbon forms on the surface. In one embodiment, the sensor is a coating of porous silicon. In other embodiments, the sensor can be a mirrored or frosted surface on the shaft. A white light source or a laser is reflected off of the sensor and an optical imaging device reads the reflected light to detect for a change in color or intensity of the light to indicate if a hydrocarbon is present on the sensor surface.

Abstract: A film cooling hole for a turbine airfoil used in a gas turbine engine, where the film cooling hole includes a metering inlet section of constant diameter cross section and a diffusion section having side walls with multiple expansion. The two side walls have an expansion of around 10 degrees and also slant inwards or toward the downstream side wall of from 10 to 45 degrees to provide the addition expansion. The downstream wall also expands at around 10 degrees. For a compound shaped film hole, the multiple expansions are 10 degrees in the downstream direction and 0 to 5 degrees in the radial outward direction. The side wall in the radial expansion direction will at a convergent angle of 10 to 45 degrees. In the radial inward direction, the expansion angle is from 10 to 20 degrees with a convergent side wall angle of from 10 to 45 degrees.

Abstract: A film cooling hole for a turbine airfoil used in a gas turbine engine, where the film cooling hole is formed from a laser with smooth surfaces and without sharp corners, the film hole having a metering section of constant diameter, a first diffusion section having a conical shape, and a spreading section having a contoured clam shell cross sectional shape that opens onto the airfoil surface. The contoured clam shell shaped spreading section includes a raised middle portion with depressions on both sides, and slanted side walls that slant toward the hole opening. The laser cut film cooling hole can be formed after the TBC has been applied.

Abstract: A turbine rotor blade made from the spar and shell construction in which the shell is a thin wall shell made from a high temperature resistant material that is formed by a wire EDM process, and where the shell is secured to the spar using a retainer that is poured into retainer occupying spaces formed in the shell and the spar, and then hardened to form a rigid retainer to secure the shell to the spar. The spar and the shell both have grooves facing each other to form a retainer groove. A retaining material, such as a liquid or a powdered metal, is poured into the grooves and hardened to form a retainer to secure the shell to the spar. The retaining material also forms a seal on the top of the spar and between the spar and shell.

Abstract: A turbine airfoil with a trailing edge cooling circuit that includes a row of pressure side bleed channels connected to a metering and impingement cooling circuit that discharges a layer of film cooling air out through a row of exit slots that open onto a pressure side wall of the trailing edge to provide cooling for the trailing edge region of the airfoil. Formed within the trailing edge tip is a vortex flow chamber extending the length of the airfoil and connected to the bleed channels by a row of inlet holes to draw a portion of the film cooling air into the vortex chamber to provide cooling for the trailing edge tip and to prevent the film cooling layer from breaking off from the pressure side wall. The vortex chamber is connected to a row of exit holes offset from the inlet holes to form the vortex flow and to discharge the vortex flowing cooling air out the trailing edge in a direction parallel to the suction side wall.

Abstract: A turbine airfoil, such as a turbine blade, having a showerhead for cooling the leading edge region. A plurality of rows of exit film slots are arranged along the leading edge in the spanwise direction and staggered from each other. The exit slots are connected to narrow cooling air passages formed within the airfoil wall in which a plurality of rows of micro pin fins extend across the passage to form serpentine flow paths for the cooling air. Cooling air from a cooling air supply channel is metered into an impingement cavity to produce backside impingement cooling of the leading edge wall. The cooling air is diffused and then passed into the plurality of narrow cooling air passages to be metered again and serpentine flow through the narrow passages, and then diffused into the exit film slots. The cooling air is then ejected from the diffusion slots as a film layer of cooling air onto the leading edge airfoil surface.

Abstract: A electrical power plant with a gas turbine engine that drives two electric high speed generators connected on the two ends of the engine rotor shaft in a direct drive relation without the need for a gear box. The high speed generators can be of the same or different power rating and connected in series such that the total power output can easily be changed without the need to redesign larger high speed generators. Because the generators use permanent magnets, they can be operated as motors to drive two or more compressors. A turbine can be used to drive two or more high speed generators connected in series. Since the high speed generator/motors have thrust bearings, the thrust bearing in the turbomachine can be eliminated.

Abstract: A turbine blade used in a gas turbine engine, the blade includes a plurality of cooled zones each with a plurality of radial extending cooling passages formed within the wall of the blade and connected to a separate collection cavity. A leading edge collection cavity is supplied with cooling air through a plurality of radial extending cooling channels located in the wall around the leading edge of the blade. Film cooling holes connected to the leading edge collection cavity discharge film cooling air to the leading edge. A pressure side collection cavity is supplied with cooling air from a plurality of pressure side radial extending cooling channels and discharges cooling air through film cooling holes on the pressure side. A suction side collection cavity is supplied with cooling air through a plurality of suction side radial cooling channels and discharges cooling air through suction side film cooling holes.

Abstract: A turbine airfoil with a leading edge cooling air supply channel located along the leading edge region of the airfoil to supply cooling air from an outside source and a series of serpentine flow cooling circuits positioned along the leading edge of the airfoil connected to the cooling supply channel to pass cooling air through the series of serpentine passages in a direction from the airfoil root to the airfoil tip. The series of serpentine circuits includes legs on the pressure side and the suction side of the leading edge. Cooling air from the supply channel is metered into the first leg of the serpentine circuit located near the root, flows through a series of serpentine circuits along the leading edge of the airfoil, and flows out to the tip through a tip hole in the last leg of the last serpentine circuit.

Abstract: A turbine rotor blade for use in an industrial gas turbine engine in which the blade includes a root section with a fillet region formed between the root and the platform of the blade. The platform region includes a diamond shaped root slot to discharge cooling air from the internal cooling passage out through the trailing edge of the blade. The diamond shaped cooling slot of the present invention comprises a smaller radius at the upper and lower corners and a much larger radius at the mid section of the slot. The cooling slot is positioned in the trailing edge root section where the upper corner is positioned above the fillet run out location. The diamond shaped cooling slot thus minimizes the high stresses induced by the stress concentration from the cooling slot and the fillet run out location.

Abstract: A turbine airfoil for use in a gas turbine engine, the airfoil formed from a support spar with a leading edge rib having a row of impingement cooling holes, the support spar having an array of modules formed on the pressure side and the suction side of the spar, and a number of cavities separated by ribs extending across the walls of the support spar. Each module is rectangular in shape and includes an impingement compartment and a diffusion compartment separated by a rib with crossover holes to connect the two compartments. Impingement holes connect the impingement compartment to a first cavity, and spent air cooling holes connect the diffusion compartment to a second cavity located downstream from the first cavity.

Abstract: A turbine blade with single tip rail located on the suction side of the tip, and a plurality of curved diffusion holes to discharge cooling air onto the pressure side wall and the tip floor or tip corner of the blade to provide cooling and to limit leakage of the hot gas flow through the tip gap. The curved diffusion holes are curved in an upward direction toward the tip, and each includes a metering hole with an inlet axis normal to the cooling supply cavity surface. A suction side tip rail is flush with the backside wall and slants outward toward the front edge on the inner side. A first curved diffusion hole opens onto the pressure side wall just below the tip corner, a second curved diffusion hole opens adjacent to the tip corner, and a third curved diffusion holes opens onto the tip floor adjacent to the tip corner and at a midpoint between the tip corner and the tip rail.

Abstract: A turbine airfoil with a multiple impingement cooling circuit to provide backside impingement cooling of the leading edge region and the pressure and suction side walls. A leading edge cavity and a mid-chord cavity are separated by a rib. A first baffle is secured within the leading edge cavity and forms a series of impingement compartments and impingement cooling holes to channel cooling air along the first baffle and provide impingement cooling along the backside wall of the leading edge region. A second baffle is secured within the mid-chord cavity and forms a series of pressure side and suction side impingement compartments and impingement cooling holes alternating from the pressure side wall to the suction side wall to channel cooling air along the second baffle and provide impingement cooling along the backside walls of the pressure and suction sides.

Abstract: A turbine rotor blade with a trailing edge cooling circuit in which cooling air from an impingement cavity located adjacent to the trailing edge region is connected to a plurality of metering channels that open onto the pressure side wall of the trailing edge region of the airfoil to discharge the cooling air. A row of submerged slots that open onto the suction side wall of the airfoil are connected to the metering channels by a row of small holes that bleed off the cooling air in the metering channel in a progressive manner and discharge the air into the slot. The trailing edge cooling circuit allows for a thinner trailing edge airfoil, reduces the metal temperature and reduces the shear mixing between the cooling air and the mainstream hot gas flow.