Abstract: A high temperature heat exchanger for use with a small gas turbine engine to produce a combined cycle power plant, where the heat exchanger includes ceramic heat exchange tubes of SiC that are tightly fitted to the heat exchanger so that no welds or brazing is used and prevent any thermal stresses between the tubes and the heat exchanger end plates or baffle plates. The heat exchanger includes an inner casing and an outer casing with the heat exchange tubes extending through the heat exchanger between the two casings, and the gas turbine engine operating in the space within the inner casing. The tubes are tightly fitted in holes within the end plates and baffle plates, and a molybdenum disulfide coating is used to form a seal.

Abstract: A turbine rotor blade with a spar and shell construction, where the shell has an airfoil shape and is formed of two shell segments with an upper shell half and a lower shell half. The upper shell half is radially supported by a tip of the spar while the lower shell half is radially loaded by an attachment so that its load is not carried by the upper shell half and the tip of the spar in order to reduce overall stress levels.

Abstract: A low pressure turbine rotor disk for a small twin spool gas turbine engine in which the rotor disk includes a forward side cavity large enough to allow for the bearing assembly that rotatably supports the rotor disk to fit within the cavity in order to shorten the axial distance between the bearings that support the inner rotor shaft on which the turbine rotor disk is secured. Minimizing the bearings spacing allows for a high critical speed for the inner rotor shaft and therefore allows for the small twin spool gas turbine engine to operate at this small scale. The turbine rotor disk also includes a plurality of axial aligned cooling air holes to allow for cooling air from the bearings to flow out from the aft end of the rotor disk. The inner surface of the cavity is an annular surface that forms a seal with knife edges extending outward from the bearing support plate also located within the cavity.

Abstract: A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of dog bone struts each mounted within openings formed within the shell and spar to allow for relative motion between the spar and shell in the airfoil chordwise direction while also forming a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from Tungsten, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: A bearingless floating wind turbine has a tall and narrow main support structure with a center of buoyancy located well above the center of gravity to provide stability to the wind turbine while supported for rotation in a body of water, a vertical axis turbine blade extends from the main support structure and rotates together under a wind, and a non-rotating shaft extends through the main support structure with a vertical axis direct drive generator connected between the main support structure and the non-rotating shaft to produce electricity when the main support structure rotates. an anchor line connected to the non-rotating shaft prevents the floating wind turbine from drifting in a body of water.

Abstract: A small single use gas turbine engine with a twin spool rotor shaft and a bypass fan, the rotor shafts being supported for rotation by bearings in the front and rear compartments of the engine. A portion of the bypass fan air is diverted to flow through the bearings in order to provide for cooling of the high speed bearings. Because a large amount of cooling air is required to cool the bearings, an air in tube is positioned in each of the guide vanes in the turbine section to provide for a cooling air passage leading into the rear bearings. In the front bearings, cooling air from the bypass fan is bled off from a location just upstream from the entrance to the centrifugal compressor. A fuel mist tube supplies fuel mist to the cooling air before passing through the bearings to provide extra cooling for the bearings and to provide lubrication to improve the service life of the engine.

Abstract: A small twin spool gas turbine engine with a bearing support arrangement in which the bearings are dampened by O-rings secured between the bearing races and the support structure, and in which the bearings are arranged in series so that a cooling air can be passed through the bearings to prevent overheating. The aft end of the engine includes high speed and low speed ball bearings supported with preload springs to add additional damping capability. The fore and aft end bearing support assemblies forms a cooling air path for the cooling fluid to flow through a passage within one of the guide vanes, through the bearings and out through a hole in the low pressure turbine rotor disk. the bearing cooling air passage includes a snorkel tube that extends from a guide vane cooling air passage and into the bypass air channel to draw in the cooling air used for the bearings. The snorkel tube includes a slanted opening so that dirt particulates do not enter the bearing cooling air passage.

Abstract: A small twin spool gas turbine engine with a bearing support arrangement in which the bearings are dampened by O-rings secured between the bearing races and the support structure, the bearings are arranged in series so that a cooling air can be passed through the bearings to prevent overheating, and the bearings are dry lubricated. The aft end of the engine includes high speed and low speed ball bearings supported with preload springs to add additional damping capability. The end bearing support assemblies form a cooling air path for the cooling fluid to flow through a passage within one of the guide vanes, through the bearings and a hole in the low pressure turbine rotor disk and out the exhaust of the engine. A snorkel tube extends into the bypass air channel drawing air for the bearings, and has a slanted opening preventing dirt particulates from entering the bearing cooling passage.

Abstract: A turbine blade for use in a gas turbine engine, where the turbine blade is made from a spar and shell construction in which a thin walled shell is formed from carbon or molybdenum nanotubes arranged in a direction such that the nanotubes are under tension when the blade is rotating in the engine. The carbon nanotubes are allotropes of carbon in which the length to diameter ratio exceeds 1,000,000 in order to produce very high tensile strength, unique electrical properties, and a very efficient conductor of heat. The nanotube shell includes a metal insert having a tear drop shape and the lower end of the shell wraps around the metal insert to form a wedge in which the shell is held in place against radial displacement between the platform and the attachment portion of the blade.

Abstract: A gas turbine engine with a turbine section having at least a first stage turbine blade and a last stage turbine blade. The first stage turbine blade includes cooling fluid passages therein in which a compressed cooling fluid, usually from the compressor section of the gas turbine engine, is passed through for cooling of the first stage blade. The last stage turbine blade includes cooling fluid passages therein, but draws the cooling air from an outside ambient pressure source instead of from a compressor. The rotation of the last stage turbine blade and rotor disk provides for a centrifugal force to drive the cooling air into the blade and through the blade for cooling thereof. No additional compression of the last stage cooling fluid is required. A cover plate with a plurality of impellers covers a back side of the last stage rotor disk and provides for an additional means to pump the ambient cooling fluid into the last stage blade.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from either Molybdenum, Niobium, alloys of Molybdenum or Niobium (Columbium), Oxide Ceramic Matrix Composite (CMC), or SiC—SiC ceramic matrix composite, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: An inner rotor shaft for use in a small twin spool gas turbine engine, the inner rotor shaft having a hollow middle section formed of a smaller diameter hollow section on a compressor end and a larger diameter hollow section on the turbine end of the shaft. Solid shaft end extend from the hollow section to form a forward solid shaft end to secure the fan rotor disk and an aft solid shaft end to secure the turbine rotor disk. A parabolic shaped transition section joins the forward shaft end to the smaller diameter hollow section, and a conical shaped transition section joins the aft shaft end to the larger diameter hollow section. A conical shaped transition piece joins the two hollow sections together to form an inner rotor shaft that can fit within a minimal space between the compressor rotor disk and the annular combustor assembly of the engine.

Abstract: A process for optimizing a blade tip clearance for a rub tolerate design in a gas turbine engine that includes evaluating a selection of candidate materials for galling and heat generation, selecting a subset of the candidate materials and analyze engine transient tip clearance for the subset materials, verifying the optimum material cooling and heat shielding for tip clearance and structures, and analyze and select an engine break-in procedure for the optimum tip clearance between turbine and compressor. If one of the subset materials results in a poor performance, than another material from the candidate materials is selected and reanalyzed until the best materials have been found. When the best materials are found, then a finite element method of analysis is performed from the blade and the static parts that form the tip clearance is performed to evaluate the materials for damage. If required, the blade and tip squealer configuration is altered to reduce stress, or blade tip coating is used.

Abstract: A process and an apparatus for generating electric current from a vibrating blade exposed to a fluid flow, in one embodiment, a plurality of blades are secured within a tail cone of a steam turbine and each blade includes a magnet and a coil associated with the magnet to produce electric current in the coil when the magnet moves across the coil. Each blade is designed to have a natural frequency so that the blade will resonate under the influence of the exhaust flow from the turbine in the tail cone. Abutment members are mounted on the blade or the fixed support to limit the range of motion of the vibrating blade so that it does not exceed the fatigue limit. The blades are made from a low damping material such as titanium so that a low frequency is produced.

Abstract: A test rig that reproduces high temperature and high pressure conditions found in a gas turbine engine for testing materials under these conditions. The test rig is formed of a plurality of modules connected in series to form the test rig. A specimen rotation module is connected to a mechanical load module, which is connected to a combustor module, which is connected to a test section module on which one or more test specimens are placed, which connects to an exhaust section module, which is connected to a thrust bearing & air inlet section module. The modules are cylindrical in shape and connected in series to form a compact test rig.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from either Molybdenum, Niobium, alloys of Molybdenum or Niobium (Columbium), Oxide Ceramic Matrix Composite (CMC), or SiC-SiC ceramic matrix composite, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: The present invention is a vane for us in a gas turbine engine, in which the vane is made of an exotic, high temperature material that is difficult to machine or cast. The vane includes a shell made from Tungsten, and is formed from a wire electric discharge process. The shell is positioned in grooves between the outer and inner shrouds, and includes a central passageway within the spar, and forms a cooling fluid passageway between the spar and the shell. Both the spar and the shell include cooling holes to carry cooling fluid from the central passageway to an outer surface of the vane for cooling. This cooling path eliminates a serpentine pathway, and therefore requires less pressure and less amounts of cooling fluid to cool the vane.

Abstract: A test facility is provided for testing materials under high temperature, pressure, and mechanical loads. The facility provides a physically sealed system that simulates conditions in hot sections or gas turbine engines. A test article is coated with a test material and exposed to a hot combusting flow in a test section housing. The article may be a pipe or conduit member oriented at any direction to the flow. A second cooler flow of fluid is channeled through the test article to create a sharp temperature gradient in the test article and through the test material. A liquid-cooled sleeve is disposed about the test article to create an annular channel of combusting flow over the test article. The downstream end of the second cooler flow is connected to the upstream end of the main hot flow at the combustion chamber.

Abstract: An inner rotor shaft for use in a small twin spool gas turbine engine, the inner rotor shaft having a hollow middle section formed of a smaller diameter hollow section on a compressor end and a larger diameter hollow section on the turbine end of the shaft. Solid shaft end extend from the hollow section to form a forward solid shaft end to secure the fan rotor disk and an aft solid shaft end to secure the turbine rotor disk. A parabolic shaped transition section joins the forward shaft end to the smaller diameter hollow section, and a conical shaped transition section joins the aft shaft end to the larger diameter hollow section. A conical shaped transition piece joins the two hollow sections together to form an inner rotor shaft that can fit within a minimal space between the compressor rotor disk and the annular combustor assembly of the engine.