Abstract: A turbine blade with a blade tip having a plurality of vortex chambers formed below the blade tip surface, the vortex chambers flowing from the suction side to the pressure side of the airfoil. Each vortex chamber includes a metering inlet hole located on the side of the chamber at the suction side end such that the cooling air entering the chamber impinges onto the backside of the blade tip and then produces a vortex flow through the chamber. An exit slot is located at the end of each vortex chamber to produce impingement cooling on the backside of the pressure side wall prior to being discharged onto the outer surface of the blade near adjacent to the pressure side corner. Helical ribs extend along the vortex chambers to promote the formation of the vortex flow.

Abstract: A rim cavity seal arrangement for a gas turbine engine, where the rim cavity is formed between two rotor disk stages with a stator vane assembly positioned between the rotor disks and extending into the rim cavity. A cover plate secured onto a side of the rotor disk includes a plurality of cooling air injectors directed to discharge cooling air into an annular groove formed on the underside of the vane endwall adjacent to the rotor disk to form an air cushion that seals the rim cavity from hot gas ingestion. The cover plate forms a cooling air supply passage with the rotor disk side to supply the injectors, and the cover plate includes at least one metering hole to meter the pressurized air into the injectors.

Abstract: A turbine rotor blade with a low cooling flow serpentine circuit to provides cooling for the airfoil. The circuit includes a three pass aft flowing serpentine circuit that begins at the airfoil mid-chord region and connects to a series of multiple impingement cooling holes formed within the trailing edge region. A double pass forward flowing serpentine circuit then connects with the triple pass aft flowing serpentine circuit to provide cooling for the leading edge region and is connected to a showerhead arrangement for discharging film cooling air. A blade tip cooling channel connects with the last leg of the double pass forward flowing serpentine to form a 6-pass serpentine flow cooling circuit for the entire blade. Since only the leading edge serpentine channel discharges film cooling air, the serpentine circuit can make use of low cooling flow to provide cooling for the entire blade.

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 turbine rim cavity seal with a variable gap clearance control in a gas turbine engine. A stator vane includes an inner shroud with an inter-stage sealing housing extending inward and separating a front rim cavity from an aft rim cavity. Front and aft rotor disks are located on the sides of the stator vane and include labyrinth seal teeth extending upward to form a seal with an abrasive material supported on the underside of the housing cavity. The abrasive material includes a slanted sealing surface. A hydraulic actuator moves the rotor disks in an axial direction to control the seal gap and regulate the purge air flow into the front rim cavity and leakage flow through the seal into the aft rim cavity. In another embodiment, the sealing surface is stepped as well as slanted.

Abstract: A turbine blade with a conical shaped tip for use in a first or second stage of a turbine. The airfoil includes an aft flowing 3-pass serpentine flow cooling circuit with a serpentine tip turn located between the first and the second legs and underneath the tip. Mini serpentine flow cooling circuits are formed in the end of the first leg and the beginning of the second leg to reduce or eliminate the cooling flow separation and over temperature issues at the tip portion of the blade. A leading edge cooling supply channel is located forward of the first leg and has a decreasing flow area with an exit cooling hole at the tip to discharge cooling air. The last leg of the main serpentine circuit is formed by a slanted rib that produces a decreasing flow area in the last leg and includes an exit cooling hole on the tip to discharge cooling air. The last leg is connected to a row of exit cooling holes or slots formed along the trailing edge of the airfoil.

Abstract: A turbine rotor blade with a low cooling flow serpentine circuit to provides cooling for the airfoil. The circuit includes a first 3-pass serpentine flow circuit with a first leg located adjacent to the leading edge to provide impingement cooling air into a leading edge impingement cavity. The remaining cooling air flows through the first serpentine circuit to provide cooling for the blade forward mid-chord region and is discharged through film cooling holes on the pressure and suction side walls. Some of the impingement cooling air for the leading edge is discharged as film cooling air for the leading edge surface while the remaining spent cooling air flows through a blade tip channel and then into the second aft flowing 3-pass serpentine circuit to provide impingement cooling for the trailing edge region before being discharged out through exit slots and blade tip corner discharge holes.

Abstract: A small twin spool gas turbine engine with a forward end of the engine supported by bearings that are cooled by passing a cooling fluid such as cooling air bled off from the compressor with a liquid lubricant sprayed into the cooling air that is then passed through the bearing to prevent overheating. The cooling fluid is then discharged from the bearing cooling circuit out the front end of the engine through a number of discharge holes formed within the threaded nut that secures the fan blade assembly to the low speed rotor shaft. The threaded nut includes a plurality of slanted holes that slant outward from the rotational axis in the direction of fluid discharge.

Abstract: A turbine rotor blade with a first cooling circuit to provide cooling for the leading edge region of the airfoil and a suction side surface of the blade tip, and a second cooling circuit to provide cooling for the mid-chord region of the blade and the pressure side surface of the blade tip. The first cooling circuit includes a cooling air supply channel connected to a leading edge impingement cavity having showerhead arrangement of film cooling holes. The cooling air supply channel is connected to the suction side tip cooling channel that extends along the tip. The second cooling circuit includes a 3-pass forward flowing serpentine circuit with the third leg connected to a pressure side tip cooling channel that extends along the pressure side edge of the blade tip. Rows of tip cooling holes connect to both of the tip channels to discharge cooling air through the blade tip sections.

Abstract: A turbine blade for use in an industrial gas turbine engine, especially for the first stage turbine blade, in which the blade includes a squealer tip with a pressure side tip rail extending all the way top the trailing edge of the tip and a suction side tip rail with a cut back adjacent to the trailing edge to form a cooling air flow path from the squealer pocket and up and over the cut back in the suction side tip rail. The floor of the squealer pocket includes a series of skewed trip strips in the location of the cut back in the tip end, the trip strips being skewed in a direction to direct the cooling air toward the cut back. Rows of film cooling holes extend along the pressure and suction sides of the blade just below the tip rails to provide cooling for the tip rails. Tip cooling holes are located on the squealer pocket floor upstream from the trip strips and along the tip rail walls to provide cooling within the pocket.

Abstract: A stator vane for a gas turbine engine having a cooling circuit that passes through the airfoil and both of the inner and the outer diameter endwalls to provide cooling for the airfoil and endwalls. Each endwall includes an intricate cooling circuit and is connected to the serpentine flow circuit within the airfoil to pass cooling air back and fourth from one endwall to the other through one of the legs in the serpentine flow circuit of the airfoil. A row of exit cooling holes is located in one or both of the endwalls and is connected to the last leg of the serpentine circuit to discharge cooling air. The airfoil and endwalls cooling circuit forms a series of passages that meander through the vane before discharging out through the exit holes.

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 includes a number of radial extending projections each with a row of cavities that form the retainer occupying spaces in order to spread the loads around. The retainer can be a bicast material, a transient liquid phase bonding material, or a sintered metal. An old shell can be easily removed and replaced with a new shell by removing parts of the retainer and re-pouring a new retainer with a new shell in place.

Abstract: A combustor for a small gas turbine engine in which a compressor discharges a swirling air flow into the combustor, and a fuel impeller is rotatably connected to the compressor and slings atomized fuel into the combustor in a direction to produce a stagnation zone within the combustor. A plurality of fuel injectors delivers fuel to the inlet end of the fuel impeller. The combustor outlet is a nozzle shape that produces a high enough pressure level in the swirl air flow such that the cooler air flows along the inner combustor wall to form an insulation layer from the hot combustor gas flow within the combustor. The fuel impeller produces a very fine mist of fuel and the stagnation zone produced results in easy starting of the engine and stable combustor burning.

Abstract: A labyrinth seal having two or more knife edges in which at least one of the knife edges has a varying radial height in the circumferential direction or a cavity formed between adjacent knife edges has a varying axial width so that the leakage flow across the gap will vary around the circumference in order that excess vibrations will not develop in the knife edges and break them. The radial height of the step that forms a gap with the lab teeth varies either sinusoidal or random. An axial spacing between the step and the associated tooth can also vary either sinusoidal or random. To minimize a cavity edge-tone effect, the corners of the steps are rounded.

Abstract: A turbine blade with a squealer pocket formed from a pressure side tip and a suction side tip rail, where two vortex cooling channels extend along the blade tip between the tip rails and an internal cooling air supply cavity to provide vortex cooling for the tip rails. Curved cooling air passages connect the cooling air supply cavity to the vortex channels and then discharge cooling air onto the pressure side wall just below the tip rail crown and into the squealer pocket from the inner surface of the suction side tip rail. Tip cooling holes are also connected to the cooling air supply cavity and discharge cooling air against the inner sides of the two tip rails to promote a vortex flow within the squealer pocket. The curved cooling holes in the vortex cooling channels are offset in order to promote the vortex flow within.

Abstract: A turbine airfoil with a film cooling hole having a bell mouth shaped opening that has expansion in both the side walls and the downstream wall of from 15 to 25 degrees. The film cooling hole includes an expansion section formed with two long ribs and one short rib to form three inlets of equal cross sectional areas so that the flows into the three passages are the same. The short rib forms two middle passages to combine with two outer passages to form four exit passages for the film hole. The two side walls are curved outward in the stream-wise oriented film hole and have an expansion of from 0 to 5 degrees in the compound angled film hole.

Abstract: A small gas turbine engine where the compressor and turbine are supported on a rotary shaft, and a main bearing is supported 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 used 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 turbine blade assembly with a spar and shell construction. The shell is held in place between a tip edge of the spar and an airfoil shaped groove on the platform surface. A segmented seal assembly is held within the groove between the shell bottom surface. The segmented seal has a top face slanted inward toward the spar to engage with a similar slanted surface formed on the bottom face of the shell. The seal segments fit within the groove to engage both side walls to form a tight seal interface between the shell bottom surface and the outer wall of the groove. The seal segments are forced upward against the slanted underside of the shell due to centrifugal forces from rotation. A cooling air passage to supply pressurized cooling air to the groove and the underside surface of the seal segments also can be used to force the seal segments up against the slanted underside of the shell.

Abstract: A large and highly twisted turbine blade for an IGT having a lower span serpentine flow cooling circuit and an upper span serpentine flow cooling circuit connected in series to provide low flow cooling for the blade. The lower span serpentine is a forward flowing 5-pass serpentine, while the upper span is an aft flowing 3-pass serpentine circuit. The last leg of the lower span serpentine and the first leg of the upper span serpentine are both aligned along the leading edge region to provide cooling there. The trailing edge includes lower span exit cooling holes and upper span exit cooling holes in which the lower span exit holes are connected to the first leg of the lower span serpentine and the upper span exit holes are connected to the last leg of the upper span serpentine. All of the cooling air from the lower span serpentine circuit that does not flow out the lower span exit holes flows into the upper span serpentine circuit to provide low flow cooling with the lower span cooled before the upper span.

Abstract: A turbine blade with a 5-pass serpentine flow cooling circuit for mid-blade cooling that is formed from a 3-pass forward flowing serpentine circuit followed by a 2-pass aft flowing serpentine circuit with the first leg or supply channel for the circuit located between the second leg and the fourth leg. A crossover channel connects the third leg to the fourth leg and passes underneath the blade tip to provide convective cooling and also discharges cooling air through tip cooling holes.