Abstract: An air cooled turbine stator vane form a small gas turbine engine in which the vane has an airfoil of less than one inch in spanwise height, where the airfoil is a hollow airfoil having an insert that forms a sequential impingement cooling circuit for the pressure side wall and then the suction side wall. The insert includes a plurality of cooling air supply channels connected together by ribs where pressure side impingement holes are formed in the cooling air supply channels and suction side wall impingement cooling holes are formed in the ribs.

Abstract: A gas turbine engine with a closed-loop liquid metal cooling fluid system for cooling stator vanes within the turbine, in which the stator vanes are made of a metallic material that will not react with the liquid metal cooling fluid. The stator vane may be made from a typical metal material such as ferrous metal alloys, nickel alloys or cobalt (Co) alloys, and an insert or liner made of molybdenum or tantalum may be placed inside to protect the outer vane material from reacting with a liquid metal such as bismuth, lead (Pb), indium, or alloy mixtures of thereof. In the case where the liquid coolant is bismuth, the liquid bismuth must be purged from the cooling system before the fluid cools and solidifies so the solidified bismuth does not expand and break the vanes.

Abstract: An air cooled turbine blade with a trailing edge cooling circuit that is formed by casting the airfoil with a pressure side trailing edge lip being oversized, and then machining away material from the pressure side wall to leave a pressure side bleed slot with a smaller (t/s) ratio than can be formed by casting alone. In another embodiment, the airfoil is cast with a trailing edge exit hole, and then the material on the pressure side wall in the trailing edge region is machined away to leave a pressure side bleed slot instead of an exit hole.

Abstract: A process for producing a small crack in a turbine rotor blade for use in training inspectors to be qualified for inspecting blades with small crack, where each blade is vibrated in order to initiate a small crack under observation using an IR camera that can detect for a small crack, and when a small crack at a desired size is produced on the blade, the vibration is stopped and the blade removed and then used in a training process to train and qualify an inspector for an inspection process.

Abstract: An extendable nozzle for a rocket engine includes a chamber with a main combustion chamber and a throat and a main nozzle formed within a guide tube, where a nozzle extension slides over the guide tube from a stowed position to an extendable position. A locking O-ring secures the nozzle extension in a stowed position, and a locking split ring holds the nozzle extension in the extended position.

Abstract: A system and a process for testing a gas turbine engine or component thereof, especially for a large aero gas turbine engine, and for a process for testing a large industrial gas turbine engine that requires large flow capacity and pressure ratios. The system and process may include the use of a large compressed air storage reservoir to provide compressed air to the testing system. Further, the system and process may also include the use of a pre-heating system, which may include a heater and a heat exchange device, to warm the compressed air from the compressed air storage reservoir to a temperature suitable to simulate normal operating conditions of the gas turbine engine or component thereof.

Abstract: An apparatus and a process for producing a shrouded centrifugal impeller where a rear hub opening is machined into the stock material so that hard to reach surfaces in the enclosed passages can be machined. When the hard to reach passages are machined to a final stage, the rear hub opening is closed using a process such as electron beam melting or electron beam welding to enclose the passages and form a single piece shrouded impeller.

Abstract: A floating air riding seal for a gas turbine engine with a rotor and a stator, an annular piston chamber with an axial moveable annular piston assembly within the annular piston chamber, an annular cavity formed on the annular piston assembly that faces a seal surface on the rotor, and a central passage connecting the annular cavity to the annular piston chamber to supply compressed air to the seal face, where the annular piston assembly is a split piston assembly to maintain a tight seal as coning of the rotor disk occurs.

Abstract: A turbine blade with a plurality of rows of micro sized near wall cooling channels formed within an outer side of the airfoil wall to produce near wall cooling of the hot surface of the airfoil. The near wall cooling channels extend around the entire airfoil surface from the trailing edge to the leading edge regions and are formed using a metal printing process. Each row of near wall cooling channels discharges into a trench that opens onto the blade tip to provide a seal against hot gas leakage across the blade tips.

Abstract: A gravity drained bearing compartment with two oil shields that are located near to an air vent and an oil scavenge line, where the oil shields function to reduce circumferential momentum of the oil to prevent oil from entering the air vent and direct oil into the scavenge line.

Abstract: A film cooling hole for an air cooled turbine airfoil, where the film cooling hole includes a first expansion section with expansion only on the downstream wall and a second expansion section with expansion on the downstream wall and the two side walls. No expansion is formed on the upstream walls on the first and second expansion sections.

Abstract: A high efficiency power plant for a UAV with a high pressure ratio gas turbine engine used for low power operation such as loiter speed and a low pressure ratio gas turbine engine used for high power operation. A power turbine receives hot gas flows from the two engines to drive an output shaft. At low power operation, only the high pressure ratio engine is operated. At high power operation, both engines are operated where the exhaust from the high pressure ratio engine is used to drive a turbine of the low pressure ratio engine. A compressor of the low pressure ratio engine supplies compressed air to a combustor that produces a hot gas stream that is passed through the power turbine.

Abstract: A method for designing an air cooled turbine airfoil includes selecting a number of different internal cooling air circuits, forming a ceramic core for each of the different cooling air circuits, forming a metal airfoil over each of the ceramic cores, leaching away the ceramic cores, mounting the airfoils in a stage of a turbine, passing a hot gas stream through the turbine, passing cooling air through each of the airfoils, and measuring a pressure and temperature differential across each of the airfoils to determine which cooling air circuit has the best performance.

Abstract: A sequential impingement cooling insert for a turbine stator vane that forms a double impingement for the pressure and suction sides of the vane or a triple impingement. The insert is formed from a sheet metal formed in a zigzag shape that forms a series of alternating impingement cooling channels with return air channels, where pressure side and suction side impingement cooling plates are secured over the zigzag shaped main piece. Another embodiment includes the insert formed from one or two blocks of material in which the impingement channels and return air channels are machined into each block.

Abstract: A large span industrial engine turbine rotor blade with a radial extending cooling air passage having a series of turbulence air mixers along the passage, where the turbulence mixers each have an inlet end and a curved and tapered surface such that cooling air is drawn into the inlet end of the mixer and then discharged from the curved and tapered surface into a middle of the passage. The cooling air flows through the passage along a series of these turbulence mixers from one mixer to another mixer along the spanwise length of the passage.

Abstract: An industrial engine turbine stator vane with a three-pass aft flowing serpentine cooling circuit and turn channels formed outside of the endwalls that connect adjacent legs of the serpentine flow circuit. A ceramic core used to cast the vane includes dual print-outs extending from the turn channel forming pieces that provide a more rigid core structure to prevent core shift or movement during casting. The dual core print-outs form purge air holes for the rim cavity that is then covered with cover plates if purge air holes are not required.

Abstract: A turbine rotor blade with a serpentine flow cooling circuit, especially a blade with a wide open tip turn, where the tip turn includes a main rib separating the two legs that are connected to the tip turn and include a bleed cooling air hole with a mini rib formed along side the main rib in the downstream leg from the tip turn in which bleed cooling air from the upstream leg flows through the hole and is impinged onto the mini rib. The bleed cooling air not only provides additional cooling for the tip turn region of the blade but eliminates the separation and recirculation issues creates in tip turns of the prior art. The bleed cooling air hole and mini rib can also be used in the root turn.

Abstract: An air cooled turbine rotor blade with a leading edge region cooling circuit having pressure and suction side feed channels separated by a rib, where the feed channels are connected to metering and diffusion cooling channels formed in the pressure and suction side walls and the middle rib that discharge into three rows of exit slots on the leading edge of the blade. The trailing edge region cooling circuit includes rows of serpentine flow circuits each having an impingement channel along a suction side and a return channel in a middle that opens into metering and diffusion channels on the pressure side that discharges into exit slots on the pressure side of the trailing edge region. The middle of the blade is cooled with a multiple pass aft flowing serpentine flow circuit with metering and diffusion film cooling slots.

Abstract: An turbine stator vane for an industrial engine, the vane having a serpentine flow cooling circuit with a first leg located along a leading edge of the vane airfoil and a partition rib separating the first leg from a second leg. A root turn channel connects the first leg to the second leg. A lower end of the partition rib includes a guide vane that directs cooling air from an aft section of the first leg toward a bleed hole for purging an ISSH. A plurality of turning guide ribs located in the aft section of the first leg also directs the cooling air toward the bleed hole.

Abstract: A rocket engine turbopump with a main rotor shaft supporting a liquid oxygen impeller on a forward end and a turbine on an aft end, and with a multiple stage liquid hydrogen impeller in-between. Two hydrostatic bearings support the main rotor shaft such that the liquid oxygen impeller and the turbine are both overhung. A balancing piston is used to balance the main rotor shaft in an axial direction. This structure allows for the main rotor shaft to be of such a large diameter that the turbopump is capable of operating at around 70,000 rpm that can produce a liquid hydrogen outlet pressure of around 4,000 psia in a turbopump having a diameter of less than 10 inches.