Abstract: A combined cycle power plant with a gas turbine engine integrated with a dirty fuel combustor in the turbine exhaust and the hot gas stream from the dirty combustor is mixed together and then passed through a heat recovery steam generator to produce steam, the steam being passed through a steam turbine to drive a second electric generator. Some of the turbine exhaust is passed directly into the HRSG while the remaining turbine exhaust is passed into the combustor and burned with the dirty fuel.

Abstract: A large and highly tapered and twisted turbine rotor blade for a large frame and heavy duty industrial gas turbine engine, where the blade includes a main spar with multiple impingement chambers extending along the chordwise direction of the blade, and with a thin thermal skin bonded to the main spar to form an airfoil section for the blade. The chordwise impingement channels are separated by ribs to form multiple chambers in the spanwise direction from the root to the blade tip. These compartmented impingement channels formed along the airfoil spanwise direction can be used for tailoring the gas side pressure variation in the spanwise direction, and individual impingement channels can be designed based on the airfoil local external heat load to achieve a desired local metal temperature. With this cooling circuit, the usage of cooling air is maximized for a given airfoil inlet gas temperature and pressure profile.

Abstract: A turbine rotor blade made from the spar and shell construction in which the shell formed from a plurality of shell segments each being a thin wall shell segment made from a high temperature resistant material that is formed by a wire EDM process, and where the shell segments are each secured to the spar separately using a retainer that is poured into retainer occupying spaces formed in the shell segments and the spar, and then hardened to form a rigid retainer to secure each shell segment to the spar individually. 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 turbine rotor blade with a tip cap section bonded to a tip end of the blade airfoil section, the tip cap section forming a separate cooling passage for the blade tip with a series of impingement chambers and diffusion chambers extending from a leading edge cooling air supply channel to a trailing edge exist hole. The chambers are formed by slanted ribs and a thin thermal skin is bonded to the ribs and upper and bottom surfaces to form an airfoil surface for the tip cap section. a blade with a damaged tip section can be repaired by removing the damaged section and bonding a tip cap section to the remaining surface of the blade to form a reusable rotor blade.

Abstract: A card seal with an annular arrangement of plates that form the card seal, where the plates are rotated at an angle around a longitudinal axis of the plates so that the gaps between adjacent plates are not parallel to the rotational axis of the rotor shaft, and where an annular disk is rotatably secured to the rotor shaft and located adjacent to card seal on the low pressure side, or on the high pressure side or even on both sides such that rotation of the annular disk pumps fluid into or draws fluid out from the plate gaps to decrease or eliminate any differential pressure across the card seal. Thus, card seal leakage is reduced or eliminated.

Abstract: An air cooled turbine airfoil, such as a rotor blade, with a sequential impingement cooling circuit that provides a high level of cooling with a low amount of cooling air. The airfoil is formed of a spar and shell construction in which the spar is formed from a series of alternating layers that form the cooling circuit with the shell forming the gap for the impingement cooling channels. Two different layers form the impingement cooling channels with a third layer being a separation layer. The impingement cooling circuit includes a first impingement channel to cooling a forward half of the pressure side wall, a second impingement cooling channel to cool the aft half of the pressure side wall, a third impingement cooling channel to cool a forward section of the suction side wall, and a fourth impingement cooling channel to cool the aft section of the suction side wall.

Abstract: An air cooled turbine blade with an array of sinusoidal shaped chordwise extending cooling air passages that start from a leading edge of the airfoil and extend to the trailing edge on both the pressure side and suction side walls. The sinusoidal shaped cooling channels on the pressure side merge with the sinusoidal shaped cooling channels on the suction side in a trailing edge region cooling channel and exit the blade through a row of trailing edge exit holes. Cooling air from a forward cooling supply cavity provides impingement cooling air for the leading edge that then flows into the sinusoidal shaped cooling channels.

Abstract: A showerhead cooling arrangement for a turbine airfoil in which the showerhead includes a row of film cooling holes on the stagnation point of the leading edge, a row of pressure side film cooling holes, and a row of suction side film cooling holes to form the showerhead. A pattern of grooves is formed on the leading edge surface in both a criss cross shape and three longitudinal shapes and in which the showerhead film cooling holes are located in the grooves. A TBC is applied over the leading edge surface and into the grooves. The grooves retain the TBC and prevent spallation, and the grooves hold the film layer together longer so that the cooling effectiveness is increased.

Abstract: A turbine rotor blade with a blade tip cooling circuit that is formed by bonding a blade tip to a top side of an airfoil. The blade tip includes a first series of ribs formed on a bottom side that defines a first layer of tip cooling holes. The top side of the airfoil includes a second series of ribs that define a second layer of cooling holes. The first and second series of ribs are formed before the blade tip is bonded to the airfoil to enclose the two layers of tip cooling holes that open onto the walls of the blade below the tip edge.

Abstract: A larger and highly twisted and tapered turbine rotor blade with a cooling circuit having a 3-pass horizontal flow serpentine circuit in a lower end of the airfoil, a 3-pass vertical flow serpentine circuit in the middle region of the airfoil, and a plurality of radial cooling channels in the upper end of the airfoil all connected in series to provide cooling for the airfoil.

Abstract: An air cooled turbine airfoil having micro cooling air passages formed within the airfoil that is of such size that the present day investment coating process cannot be used. The internal cooling air features of the airfoil are formed using the Tomo Lithographic Molding (TLM) process. The TLM process is a low pressure casting process that can produce precise micro-features integral to macro-scale structures using any of the exotic alloys or other metallic materials currently being used in airfoil production. The normal internal cooling air passages as well as very small features such as trip strips, pin fins, dimples, pedestals and enclosed passages such as film holes can be produced using the TLM process.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A transition duct for a gas turbine engine of the kind with can combustors, the transition duct is formed from an inner section and an outer section that form a spiral shaped cooling channel from an inlet end to an outlet end. The spiral channel includes a number of fins that extend into the cooling air passage. an annular plate with an annular arrangement of exit holes closes an aft end of the spiral cooling channel and forms an expansion chamber for the exit holes.

Abstract: A turbine blade having a squealer tip rail forming a squealer pocket on the blade tip, and a row of blade tip peripheral film cooling holes on the pressure side and suction side of the blade for cooling the blade tip rails. The pressure side tip peripheral holes extend only along a mid chord region and the suction side peripheral holes extend along the leading edge region only. A TBC is applied to the pressure side and suction side walls of the blade up to the row of tip peripheral film cooling holes, leaving these surfaces uncovered. The squealer pocket is covered with TBC while the top surfaces of the tip rails are uncovered. The surface of the airfoil above the row of tip peripheral cooling holes is without a TBC so that the metal surface will be exposed to the layer of film cooling holes discharged from the tip peripheral cooling holes.

Abstract: A turbine airfoil includes a plurality of mini serpentine flow passages to provide near wall cooling to a thin thermal skin that forms the outer airfoil surface. The airfoil includes a spar with a number spanwise extending external ribs with a series of mini pin fins extending between adjacent ribs. The thin skin includes an arrangement of mini pin fins that, when bonded to the spar, form the serpentine flow passages. The thin thermal skin is bonded to the spar to form the airfoil. The mini serpentine flow passages can be formed in the airfoil walls with a small size that cannot be formed in an airfoil formed by the investment casting process. With the use of a thin thermal skin, higher levels of cooling can be produced using the near wall cooling process.

Abstract: A turbine airfoil with pressure and suction side walls having a series of radial extending multiple impingement cooling channels in which each channel is formed with a series of impingement chambers and impingement holes that discharge impingement cooling air against the backside walls of the airfoil. The spent impingement cooling air from the radial impingement channels is discharged into collector cavities and then discharge as film cooling air onto the external surface of the airfoil.

Abstract: A cryogenic turbo-pump with an inducer and a system for detecting an onset of cavitation at the blade tips of the inducer. An IR imaging sensor is used to view the blade tips during operation of the turbo-pump for a temperature drop that indicates the onset of cavitation. The imaging sensor includes a means to vary the image capture rate to that the rotating blade tip will appear to be stationary in order to measure the blade tip temperature. A window is formed in the inducer housing in which the imaging sensor can view the blade tips during operation.

Abstract: A turbine blade with a low flow cooling circuit that includes two 5-pass serpentine flow circuits that are partially separated and partial combined to form the low flow capability while providing adequate cooling for the blade. The pressure sidewall and the suction sidewall both include an up-pass channel and a down-pass channel to form the first two legs of two serpentine flow circuits. Positioned between the up-pass and down-pass channels are two mid-chord channels that form third and fourth legs of the common serpentine flow circuit. A fifth leg is formed through a trailing edge up-pass channel that provides cooling air for a trailing edge cooling circuit with exit holes. The forward most mid-chord chamber that forms the third leg supplies impingement cooling air to the leading edge cooling circuit that also includes film cooling holes for the leading edge surface.

Abstract: A blade outer air seal segment assembly for a turbine in which adjacent segments form a curved metering and diffusion cooling slot for the discharge of film cooling air onto an inner surface of the segments. The segments include a first diffusion cavity connected to the backside surfaces of the segments through metering holes. The curved slot is connected to the first diffusion cavity. spend cooling air for cooling of the backside surfaces of the segments is metered into the first diffusion cavity and then discharged as film cooling air from the curved slot in a direction of rotation of the rotor blades to provide both sealing and cooling for the mate face of the segments.

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 includes a combustor surrounded by an upper plenum chamber and an exhaust plenum that receives a hot gas stream from the combustor. A transparent channeling vessel guides the hot gas stream from the combustor over a test specimen on which a material to be tested is mounted. The hot gas stream exits the clear channeling vessel and into the exhaust plenum where the stream is cooled by diluting the stream with cooling air. A viewing portal is located in the exhaust plenum so that the material can be observed through the clear vessel.