Patent application number | Description | Published |
20080281744 | SYSTEM AND METHOD FOR DETERMINING THE LIQUIDITY OF A CREDIT - The present invention relates to a credit index, a system and method for structuring a credit index, a system and method for operating a credit index, and a system and method for determining the liquidity of a credit. | 11-13-2008 |
20080313097 | CREDIT INDEX, A SYSTEM AND METHOD FOR STRUCTURING A CREDIT INDEX, AND A SYSTEM AND METHOD FOR OPERATING A CREDIT INDEX - The present invention relates to a credit index, a system and method for structuring a credit index, a system and method for operating a credit index, and a system and method for determining the liquidity of a credit. | 12-18-2008 |
20120095939 | SYSTEM AND METHOD FOR DETERMINING THE LIQUIDITY OF A CREDIT - The present invention relates to a credit index, a system and method for structuring a credit index, a system and method for operating a credit index, and a system and method for determining the liquidity of a credit. | 04-19-2012 |
20140012778 | SYSTEM AND METHOD FOR DETERMINING THE LIQUIDITY OF A CREDIT - The present invention relates to a credit index, a system and method for structuring a credit index, a system and method for operating a credit index, and a system and method for determining the liquidity of a credit. | 01-09-2014 |
Patent application number | Description | Published |
20090104035 | ADVANCED TURBULATOR ARRANGEMENTS FOR MICROCIRCUITS - A passageway is provided through which a cooling fluid flows in a first direction. The passageway has a plurality of trip strips positioned within the passageway. Adjacent one of the trip strips are oriented to converge towards each other at a first end to form an apex portion and to form a region in which turbulence is created. The apex portion is oriented at an angle with respect to the first direction. | 04-23-2009 |
20090116956 | MANUFACTURABLE AND INSPECTABLE COOLING MICROCIRCUITS FOR BLADE-OUTER-AIR-SEALS - A method for manufacturing a cooling microcircuit in a blade-outer-air-seal is provided. The method broadly comprises the steps of forming a first section of the blade-outer-air-seal having a first exposed internal wall, forming a second section of the blade-outer-air-seal having a second exposed internal wall, and forming at least one cooling microcircuit on at least one of the first and second exposed internal walls. | 05-07-2009 |
20100158669 | MICROCIRCUITS FOR SMALL ENGINES - A turbine engine component for use in a small engine application has an airfoil portion having a root portion, a tip portion, a suction side wall, and a pressure side wall. The suction side wall and the pressure side wall have the same thickness. Still further, the turbine engine component has a platform with an internal cooling circuit. | 06-24-2010 |
20100221098 | Peripheral Microcircuit Serpentine Cooling for Turbine Airfoils - A turbine component has an airfoil portion with at least one central core element, a pressure side wall, and a suction side wall. The airfoil portion also has a serpentine cooling passageway in at least one of the walls. In a preferred embodiment, the airfoil portion has a serpentine cooling passageway in both of the pressure and suction side walls. A refractory metal core for forming the serpentine cooling passageway(s) is also described. | 09-02-2010 |
20120152484 | PERIPHERAL MICROCIRCUIT SERPENTINE COOLING FOR TURBINE AIRFOILS - A turbine component has an airfoil portion with at least one central core element, a pressure side wall, and a suction side wall. The airfoil portion also has a serpentine cooling passageway in at least one of the walls. In a preferred embodiment, the airfoil portion has a serpentine cooling passageway in both of the pressure and suction side walls. A refractory metal core for forming the serpentine cooling passageway(s) is also described. | 06-21-2012 |
Patent application number | Description | Published |
20090304516 | PLATFORM MATE FACE CONTOURS FOR TURBINE AIRFOILS - Gas turbine engine components having an airfoil extending outwardly of a platform are mounted in adjacent relationship, and such that cooling air flows outwardly of a gap between mating faces of the platforms. The location of localized hot spots is identified on the platform, and the mating faces are designed to provide cooling air through the gap to address these hot spots. A suction side edge of the platform has a curved portion extending inwardly into the platform, and the pressure side has a curved portion bulging outwardly away from the airfoil. When these two portions on adjacent components mate, a gap is provided between two platforms that provides leakage cooling air to the hot spot. | 12-10-2009 |
20130019604 | MULTI-STAGE AMPLIFICATION VORTEX MIXTURE FOR GAS TURBINE ENGINE COMBUSTOR - A multi-stage vortex mixer for a combustor of a gas turbine engine includes a vortex amplifier stage in communication with a first stage amplifier, the vortex amplifier stage in communication with a dilution hole. | 01-24-2013 |
20130025287 | DISTRIBUTED COOLING FOR GAS TURBINE ENGINE COMBUSTOR - A combustor component of a gas turbine engine includes a refractory metal core (RMC) microcircuit for self-regulating a cooling flow. | 01-31-2013 |
20130025288 | MICROCIRCUIT COOLING FOR GAS TURBINE ENGINE COMBUSTOR - A combustor component of a gas turbine engine includes a refractory metal core (RMC) microcircuit. | 01-31-2013 |
20130327048 | COMBUSTOR LINER WITH CONVERGENT COOLING CHANNEL - A combustor liner includes a heat shield, a shell, a series of trip strips, and a series of projecting walls. The heat shield has a shield cold side. The shell is attached to the heat shield and includes a shell hot side facing the shield cold side, a shell cold side facing away from the shield cold side, and a row of cooling holes. The trip strips run parallel to each other and all project from the shield cold side the same distance. Each projecting wall runs parallel to, and opposite of, a corresponding trip strip and projects from the shell hot side such that the distance to which each projecting wall projects is greater for projecting walls farther from the row of cooling holes, creating successive gaps between the projecting walls and corresponding trip strips that decrease from the row of cooling holes to create a convergent channel. | 12-12-2013 |
20130327049 | COMBUSTOR LINER WITH REDUCED COOLING DILUTION OPENINGS - A combustor liner is arcuate in shape and defines an axis and a circumferential direction. The combustor liner includes a first row of dilution openings and a second row of dilution openings. The first row runs in the circumferential direction. The second row runs parallel to the first row and is axially spaced from the first row. Each dilution opening of the second row overlaps in an axial direction a portion of each of two adjacent dilution openings of the first row. | 12-12-2013 |
20130327056 | COMBUSTOR LINER WITH DECREASED LINER COOLING - A shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall. | 12-12-2013 |
20130327057 | COMBUSTOR LINER WITH IMPROVED FILM COOLING - A heat shield for a combustor liner includes first linear film cooling slots through the heat shield and second linear film cooling slots through the heat shield. The first linear film cooling slots are run in a row and each of the first linear film cooling slots is angled from the row in a first direction. The second linear film cooling slots also run in the row and each of the second linear film cooling slots is angled from the row in a second direction opposite the first direction. The second linear film cooling slots alternate with the first linear film cooling slots in the row. The first and second linear film cooling slots are connected to form a single, multi-cornered film cooling slot. | 12-12-2013 |
20140083100 | GAS TURBINE ENGINE COMBUSTOR - A swirler assembly for a gas turbine engine includes an outer annular injector which at least partially surrounds an inner injector. | 03-27-2014 |
20140090402 | Combustor Bulkhead Assembly - A heat shield is disclosed. The heat shield may comprise a body having a back surface and an opposite front surface, wherein an opening in the body communicates through the front and back surfaces. The heat shield may further comprise at least one radial rail disposed on the back surface and extending radially outward from the opening for directing cooling air flow. | 04-03-2014 |
20140096528 | Cooling for Combustor Liners with Accelerating Channels - A combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like. | 04-10-2014 |
20140137559 | GAS TURBINE ENGINE COMBUSTOR WITH INTEGRATED COMBUSTOR VANE - A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber. | 05-22-2014 |
20140216044 | GAS TURBINE ENGINE COMBUSTOR HEAT SHIELD WITH INCREASED FILM COOLING EFFECTIVENESS - A heat shield for a gas turbine engine includes a hot side with one or more raised features that extend therefrom. | 08-07-2014 |
20140338336 | GAS TURBINE ENGINE COMBUSTOR WITH INTEGRATED COMBUSTOR VANE - A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber. | 11-20-2014 |