Abstract: A blade tip shroud segment for a gas turbine engine in which the shroud segment is cooled and sealed using a combination of vortex cooling air flow and cooling air discharge film cooling. The shroud segment includes a row of vortex chambers formed within the shroud segment and a row of circumferential circular grooves formed on the hot gas flow surface of the shroud segment, where each groove is connected to a vortex chamber through a slot. Cooling air is to the vortex chambers from individual cooling air feed slots or through inter-linking channels connecting adjacent vortex chambers.

Abstract: A turbine blade for use in a gas turbine engine that requires internal cooling for the blade. The blade includes a leading edge cooling supply cavity and a trailing edge cooling supply cavity to supply pressurized cooling air from an external source to the internal cooling circuit of the blade. The leading edge cooling supply cavity is connected to a leading edge impingement cavity through metering and impingement holes to provide metering and impingement cooling air to the leading edge region of the blade and to the showerhead film cooling holes. Both the leading edge and the trailing edge cooling supply cavities also supply cooling air through metering and impingement holes to respective pressure side and suction side impingement cavities located between the two supply cavities for impingement cooling. Each impingement cavity is connected to the blade external surface through a row of film cooling holes to discharge the spent impingement cooling air.

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 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.

Abstract: An industrial gas turbine engine capable of operating at higher temperatures than air cooled nickel based alloy airfoils in the turbine. The engine burns stoichiometric or rich to produce a hot gas stream that is mostly without oxygen, and the turbine airfoils are made from a high temperature resistant material such as a refractory material that also has poor oxidation resistance. A second small gas turbine engine is used to compress nitrogen gas that is used to pass through the turbine airfoils for cooling where the film cooling nitrogen gas is injected into the hot gas stream but without igniting the fuel rich gas stream.

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 blade with a leading edge cooling circuit formed from a series of impingement cavities spaced along the leading edge and separated from the serpentine flow cooling circuit in the remaining sections of the airfoil in order to provide for a low cooling flow along the leading edge. The multiple impingement cavities are separated by a slanted rib on the bottom of the cavity, and each slanted rib includes an impingement hole directed to discharge impinging cooling air to the backside wall of the leading edge. Cooling air flows through the series of impingement cavities and into the last impingement cavity formed at the blade tip. The cooling air then is discharged through tip exit cooling holes. The blade with the multiple impingement cavities formed along the leading edge region can be cast using the prior art casting techniques.

Abstract: A stepped labyrinth seal having a plurality of teeth forming gaps between a stationary honeycomb member, where the cavities formed between adjacent teeth include a plurality of vortex flow generating strips on the wall of the cavity that redirect the leakage flow from a gap back toward the gap and in a direction toward the gap in order to limit the leakage flow through the gap. The vortex generating strips extend along the cavity floor and up the sides of the teeth forming the cavity. The strips slope downward from the downstream tooth toward the upstream tooth in the direction of rotation, and with a curvature in the direction of rotation. The stepped labyrinth seal is intended for use in a gas turbine engine having a plurality of stages in the turbine, with the labyrinth seal being used to seal the gap between rotating blades and the honeycomb surface on the stationary casing.

Abstract: A turbine blade for a gas turbine engine, in which the turbine blade includes an airfoil portion with a root having a dovetail shape, and two platform halves that include a dovetail shaped opening within the platform halves to secure the blade root within the platform halves when fastened together. The platform halves have an outer fir tree shaped surface so that the blade assembly can be inserted into a slot within a rotor disk. the blade is uncoupled from the platform in the invention so that the airfoil can be made from a single crystal material with low casting defects because the platform is not cast with the airfoil. the two platform halves include the openings with side walls that are curved to follow the contour of the airfoil root so that the airfoil is secured within the platform halves against all directions of movement.

Abstract: An apparatus and process for testing airflow through a turbine vane under actual engine airflow conditions. The apparatus includes first a second static pressure chamber each with an opening on which a cover plate is to be secured to enclose the chamber and hold the vane in place for testing. Each chamber includes a pressurized air supply tube that extends into the chamber to deliver the volume of air necessary for testing the airflow. The chamber is of such a size that the high volume of pressurized air delivered to the chamber will enter the vane as static air for the testing. The cover plates are formed of two halves in which the vane endwall can be secured between the halves. A stand supports both cover plates so that the plates and the vane can be moved into position where the cover plates are secured to the opening of the chamber for testing. The pressurized air supplied to both chambers flows into the vane and through the film cooling holes so that proper testing of the airflow can be done.

Abstract: A turbine rotor blade made from a spar and a shell each having an airfoil cross sectional shape and each formed from a high temperature exotic material and formed by a wire EDM process to form a thin wall shell from Molybdenum and the spar from Waspalloy or IN100 in order to form a turbine blade that requires very low amounts of cooling air while allowing for high temperature use above what nickel super alloys can be exposed to. The spar and shell are secured to the platform by a bolt passing through the widest section of the hollow spar and connected at one end to a tip cap and to the other end to the platform and root piece of the blade.

Abstract: A second stage turbine blade for an IGT includes a leading edge cooling supply channel to provide convection cooling to the leading edge region, and a blade tip cooling channel connected downstream from the leading edge supply channel to provide cooling for the blade tip region. A three-pass aft flowing serpentine cooling circuit provides convection cooling for the mid-chord region and discharges cooling air through a row of trailing edge exit holes or slots. A re-supply hole connects the end of the tip cooling channel to the second leg of the serpentine flow circuit to merge the tip channel cooling air with the serpentine flow cooling air before being discharged through the exit holes.

Abstract: A composite turbine blade for use in a gas turbine engine under high gas flow temperature. The blade includes a root and platform portion formed out of the standard high temperature resistant metallic material of the prior art blades and with a near wall cooled metallic spar extending along the suction side wall of the blade and wrapping around the leading edge and the trailing edge portions of the blade. The wrap around spar forms a concave opening in which a ceramic or carbon fiber mid chord section of the blade is secured to form the composite blade. The solid ceramic or carbon fiber mid-chord blade is exposed to the pressure side hot gas flow and requires no cooling. Radial cooling channels with pin fins extend along the blade spar on the leading and trailing edges and on the suction side wall to provide cooling for the metallic portions of the composite blade.

Abstract: A thermal mechanical fatigue test rig for testing a coating, such as a thermal barrier coating, under high temperature and pressure to simulate the actual operating environment of the coating. The test rig includes a combustor to produce a hot gas flow, a hollow test specimen on which the coating is placed, and a sapphire vessel that encloses the hollow test specimen to form a hot gas flow path over the coating. The sapphire vessel is clear so that the coating can be observed by a camera during the testing. An exhaust plenum is formed around the sapphire vessel to collect the exhaust form the hot gas flow in which additional cooling air and water for quenching can be injected to reduce the temperature of the hot gas flow prior to discharge from the test rig.

Abstract: A turbine blade, especially a turbine blade for the first or second stage of an industrial gas turbine engine, the turbine blade having an aft flowing triple pass serpentine cooling circuit with all convection cooled blade. The three passes or legs of the serpentine flow circuit are formed by a leading edge rib and a trailing edge rib that are both slanted in order to provide decreasing flow cross sectional areas in the three passes of legs. A mini serpentine flow circuit is formed on the blade tip and the beginning of the second leg of the circuit to prevent flow separation of the cooling air at the blade tip. The third leg of the circuit includes pin fins to promote heat transfer. A row of exit cooling slots are located along the trailing edge of the blade and are connected to the third leg to discharge cooling air out from the trailing edge region. Trip strips are located on the walls of the first and second legs to also promote heat transfer within these passages.

Abstract: A turbine blade with a dual serpentine flow cooling circuit to provide cooling to the airfoil section and to the blade tip section. An aft flowing multiple pass serpentine circuit includes a first leg located adjacent to the leading edge region of the blade and is connected to a leading edge impingement cavity through a row of impingement holes. A showerhead arrangement is connected to the impingement cavity to provide film cooling to the leading edge of the blade. A second aft flowing serpentine circuit is located in the tip region and is connected to the impingement cavity to supply the cooling air to the second serpentine. A row of impingement holes and exit cooling slots is arranged along the trailing edge region and connected to the first serpentine. A trailing edge discharge hole is located in the tip region to discharge cooling air from the second serpentine. The second serpentine includes film cooling holes on the pressure side to discharge film cooling air.

Abstract: A small twin spool gas turbine engine with a hollow inner rotor shaft having solid shaft ends and an outer rotor shaft having a cylindrical portion on the compressor end that forms a forward bearing support surface and a turbine rotor disk on the turbine end that forms a aft bearing support surface. The inner rotor shaft includes solid shaft ends that project out from the cylindrical portion of the outer shaft on one end and out from the turbine rotor disk on the other end. An inner bearing housing is secured on the solid shaft ends of the inner rotor shaft. A threaded nut on the inner rotor shafts ends provide a compressive load to the inner bearing housings which results in a tension preload to the inner rotor shaft solid ends so that the bearing assemblies for the forward and aft ends of the twin spools do not become lose from the engine operation.

Abstract: A small gas turbine engine with a compressor, a combustor, and a turbine located downstream of the combustor. The compressor and turbine are supported on a rotary shaft, and a main bearing is support on the rotary shaft, the main bearing being located in a hot zone of the combustor. The main bearing includes cooling air passages within the races to provide cooling for the bearing. A cooling air is diverted from the compressor and passed through the bearing cooling passages for cooling the bearing, and then the cooling air is directed into the combustor. The cooling air is also passed through a guide nozzle before being passed through the bearing to cool both the guide nozzle and the bearing. A swirl cup injector is sued to deliver the compressed air from the compressor and the cooling air from the bearing into the combustor, the swirl cup injector also acting to draw the cooling air through the bearing.

Abstract: A process for determining a fluid flow velocity and then a mass flow rate of a cryogenic fluid in which a Helmholtz resonator is used to detect a presence of a two-phase flow in a conduit. The process first measures a frequency of the fluid flow, then determines a speed of sound of the fluid from the frequency, then measures the temperature of the fluid flow, and from the speed of sound and the temperature determines a quality of the fluid flow (liquid or vapor), and from the frequency and the fluid quality determines a Strouhal Number. The fluid flow velocity is found from an equation relating the frequency and a diameter of the surface area to the Strouhal Number. The mass flow rate is found from the fluid flow velocity and the temperature.

Abstract: A small gas turbine engine with a bearing cooling and lubricating passage arrangement for a high speed rotor shaft. The engine includes a bypass fan and a compressor. The rotor shaft includes a central passage extending through the entire shaft, and where the rotor shaft is supported by a forward bearing and a rearward bearing. Cooling air for the bearings is diverted from the bypass air and is channeled through the bearings. Fuel is added to the cooling air at a location upstream of the bearing to provide lubrication. The cooling air and lubricating fuel passes through the bearings and into the rotating central shaft, and is then forced to flow toward a radial passage located adjacent to the combustor. The fuel is collected on the central shaft surface and forced out the radial passage and into the combustor. The cooling air continuous out from the central shaft to be mixed with the engine exhaust.