Patent application number | Description | Published |
20130019898 | EFFLUENT COLLECTION UNIT FOR ENGINE WASHING - An effluent collection unit for engine washing is formed by a portable trailer having a plurality of sides forming an internal compartment and an effluent collection system positioned within the internal compartment, which effluent collection device captures engine wash water effluent when the trailer is placed in an operational configuration. | 01-24-2013 |
20130023378 | Method of Assembly for Gas Turbine Fan Drive Gear System - A method of assembling an epicyclic gear train includes the steps of providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced apart apertures provided at an outer circumference of the carrier. Gear pockets are provided between the walls and extend to the apertures, and a central opening in at least one of the walls. A plurality of intermediate gears is inserted through the central opening and move the intermediate gears radially outwardly into the gear pockets to extend through the apertures. A sun gear is inserted through the central opening. The plurality of intermediate gears is moved radially inwardly to engage the sun gear. A gear reduction is also disclosed. | 01-24-2013 |
20130037617 | RFID TAG SYSTEM - An identification system for machinery or equipment, such as an aircraft engine, includes a name plate, a mounting bracket for mounting the name plate, a housing, and an RFID tag positioned within the housing between the name plate and the mounting bracket. The RFID tag includes an integrated circuit and an antenna such as a patch antenna. | 02-14-2013 |
20130047579 | GAS TURBINE ENGINE FAN VARIABLE AREA NOZZLE WITH SWIVABLE INSERT SYSTEM - An example nacelle assembly for a gas turbine engine includes a nacelle defined about an axis and defining a boundary of a fan bypass flow path. A fan variable area nozzle includes a plurality of inserts movably mounted to said nacelle. Each of the multiple of inserts is located at a circumferential position about the nacelle. The multiple of inserts are each independently moveable into the fan bypass flow path relative the nacelle to selectively vary a fan nozzle exit area. | 02-28-2013 |
20130065048 | LAYERED THERMAL BARRIER COATING WITH BLENDED TRANSITION AND METHOD OF APPLICATION - A multilayer coating includes a bond coat layer, a first barrier layer applied on the bond coat layer, and a second barrier layer applied on the first barrier layer. The first barrier layer has a compositional gradient comprising a majority of a first rare earth stabilized zirconia material proximate the bond coat layer to a majority of a second rare earth stabilized zirconia material away from the bond coat layer. The first and second rare earth stabilized zirconia materials are different. The second barrier layer has a compositional gradient comprising a majority of the second rare earth stabilized zirconia material to 100 wt % of a third rare earth stabilized zirconia material away from the first barrier layer. | 03-14-2013 |
20130067717 | METHOD OF INSTALLING A FLEXIBLE COMPONENT IN A U-SHAPED COMPONENT - An installation method makes use of a tool that includes a block and a plate. The block is configured to align the installation tool with a U-shaped component. The plate extends from the block and is configured to hold a flexible component, such that the flexible component is accurately located within the U-shaped component. | 03-21-2013 |
20130068729 | Method and Apparatus for Microplasma Spray Coating A Portion of a Compressor Blade in a Gas Turbine Engine - A method and apparatus for microplasma spray coating a portion of a substrate, such as a gas turbine compressor blade, without masking any portions thereof. The apparatus includes a microplasma gun with an anode, cathode, and an arc generator for generating an electric arc between the anode and cathode. An arc gas emitter injects inert gas through the electric arc. The electric arc is operable for ionizing the gas to create a plasma gas stream. A powder injector injects powdered material into a plasma stream. A localized area of the compressor blade is coated with the powdered material without having to mask the compressor blade. | 03-21-2013 |
20130092340 | Process and Refractory Metal Core for Creating Varying Thickness Microcircuits for Turbine Engine Components - The present disclosure is directed to a refractory metal core for use in forming varying thickness microcircuits in turbine engine components, a process for forming the refractory metal core, and a process for forming the turbine engine components. The refractory metal core is used in the casting of a turbine engine component. The core is formed by a sheet of refractory metal material having a curved trailing edge portion integrally formed with a leading edge portion. | 04-18-2013 |
20130094946 | TURBINE SHROUD THERMAL DISTORTION CONTROL - A shroud for a gas turbine engine includes a leading portion having a leading edge and a first set of circumferentially spaced slots at the leading edge that break up the leading portion into circumferentially spaced segments separated by the first set of slots, and a trailing portion adjacent to the leading portion. The trailing portion has a trailing edge. | 04-18-2013 |
20130098879 | Method and Apparatus for Microplasma Spray Coating a Portion of a Turbine Vane in a Gas Turbine Engine - A method and apparatus for microplasma spray coating a portion of a turbine vane without masking any portions thereof. The apparatus includes a microplasma gun with an anode, cathode, and an arc generator for generating an electric arc between the anode and cathode. An arc gas emitter injects gas through the electric arc. The electric arc is operable for ionizing the gas to create a plasma gas stream. A powder injector injects powdered material into a plasma stream. A localized area of the turbine vane is coated with the powdered material without having to mask the turbine vane. | 04-25-2013 |
20130101737 | Non-Stick Masking Fixtures and Methods of Preparing Same - Non-stick fixtures for selectively masking portions of a workpiece during application of a workpiece coating are described herein. These fixtures have predetermined surfaces thereon having an average surface roughness of about 25 Ra or less and a Rockwell hardness of about 65 Rc or more. The controlled average surface roughness ensures that these fixtures are non-stick with respect to the workpiece coating being applied to the workpieces disposed therein. The controlled Rockwell hardness ensures that the desired average surface roughness can be maintained throughout repeated use of the fixture in harsh coating environments. These fixtures reduce the workpiece coating bridging that occurs between the fixture and the workpiece, and also reduce the amount of overspray that occurs on the workpiece, thereby minimizing the amount of handwork and/or rework that is necessary after the workpiece is coated. This improves process cycle times and yields significantly. | 04-25-2013 |
20130121824 | GAS TURBINE ENGINE HAVING CORE AUXILIARY DUCT PASSAGE - A gas turbine engine system according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly defined about an axis and a core engine positioned radially inward of the nacelle assembly and having a core passage and at least one core auxiliary duct passage. The at least one core auxiliary duct passage includes an inlet for receiving a portion of a core airflow from the core engine and an outlet for discharging the portion of the core airflow. At least one of the inlet and the outlet are selectively translatable to divert the portion of the core airflow into the at least one core auxiliary duct passage and a mixer disposed between the nacelle assembly and the core engine. | 05-16-2013 |
20130142629 | Preforms and Related Methods for Repairing Abradable Seals of Gas Turbine Engines - A preform, for repairing an abradable seal component of a gas turbine engine, includes a multilayer stack that has a first layer and a second layer. The first layer is operative to bond the multilayer stack to a structural substrate of an abradable seal. The first layer includes structural material that corresponds to a material of the structural substrate and braze material that is compatible with the structural material. The second layer includes an abradable material. The multilayer stack is operative to bond to the structural substrate of the abradable seal component during a brazing process such that the second layer forms a replacement abradable layer of the abradable seal component. | 06-06-2013 |
20130146266 | SHOT TUBE PLUNGER FOR A DIE CASTING SYSTEM - A method for controlling a temperature of a portion of a die casting system having a shot tube plunger, according to an exemplary aspect of the present disclosure includes, among other things, communicating a fluid through a fluid inlet of a fluid passageway of a thermal control scheme of the shot tube plunger. The fluid circulates through the fluid passageway of the thermal control scheme to selectively adjust a temperature of the shot tube plunger. The fluid is discharged through a fluid outlet of the fluid passageway. | 06-13-2013 |
20130164137 | TURBOFAN FLOW PATH TRENCHES - An integrally bladed disk includes a rotor disk and circumferentially spaced first and second blades. The rotor disk has a rim the periphery of which forms a flow surface. The first and second blades extend integrally outward from the rim. The rim defines a trench in the flow surface between the first and second blades aft of a leading edge of the rim. The trench extends axially forward and rearward of a leading edge of the first blade. | 06-27-2013 |
20130170950 | VARIABLE SHAPE INLET SECTION FOR A NACELLE ASSEMBLY OF A GAS TURBINE ENGINE - A method of improving the aerodynamic performance of an inlet section of a nacelle assembly of a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, (a) sensing an operability condition, (b) adjusting a leading edge of the inlet section in response to the step (a), and (c) adjusting a thickness of the inlet section in response to the step (a). | 07-04-2013 |
20130170958 | GAS TURBINE ENGINE FRONT CENTER BODY ARCHITECTURE - A gas turbine engine includes a central body support that provides an inner annular wall for a core flow path. The central body support includes a first mount feature. A geared architecture interconnects a spool and a fan rotatable about an axis. A flex support interconnects the geared architecture to the central body support. The flex support includes a second mount feature that cooperates with the first mount feature for transferring torque there between. A method of disassembling a front architecture of a gas turbine engine includes the steps of accessing forward-facing fasteners that secure a central body support to a flex support, wherein the flex support includes a geared architecture supported thereon, removing the fasteners, and decoupling first and second mount features respectively provided on the central body support and the flex support. | 07-04-2013 |
20130175329 | DIFFUSION BONDING MACHINE AND METHOD - An example diffusion bonding machine includes a support structure configured to receive first and second die sets. A heat transfer device is arranged near the support structure and is configured to transfer heat relative to the die sets. A mechanism is configured to separate the die sets from one another during heat transfer. In one example method of diffusion bonding, heat is transferred relative to a space between die sets. The die sets are supported on the support structure, and a load is applied to the die sets to diffusion bond a component within each of the die sets. | 07-11-2013 |
20130186099 | Gas Turbine Combustor - A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine. | 07-25-2013 |
20130192199 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft. | 08-01-2013 |
20130192200 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 08-01-2013 |
20130192201 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 08-01-2013 |
20130192258 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 08-01-2013 |
20130192266 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The second turbine is disposed forward of the fan drive turbine. The fan drive turbine includes at least three rotors and at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between about 2.00 and about 2.30. A speed change system is driven by the fan drive turbine for rotating the fan about the axis. | 08-01-2013 |
20130192268 | INTERNALLY COOLED SPOKE - A turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, a high pressure turbine in fluid communication with the combustor, a low pressure turbine in fluid communication with the high pressure turbine, and a mid turbine frame located axially between the high pressure turbine and the low pressure turbine. The mid turbine frame includes an outer frame case, an inner frame case, and a plurality of hollow spokes that distribute loads from the inner frame case to the outer frame case. The spokes are hollow to allow cooling airflow to be supplied through the spokes to the inner frame case. | 08-01-2013 |
20130192996 | SURFACE IMPLANTATION FOR CORROSION PROTECTION OF ALUMINUM COMPONENTS - An aluminum alloy component has a surface region alloyed with an anodic metal to increase corrosion resistance in aqueous environments with high salinity or sulfur content. | 08-01-2013 |
20130199200 | FUEL DISTRIBUTION WITHIN A GAS TURBINE ENGINE COMBUSTOR - A fuel system for a gas turbine engine includes a plurality of duplex nozzles arranged on each side of top dead center and a plurality of simplex nozzles. A primary manifold is operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles and a secondary manifold is operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. An equalizer valve that is in communication with both the primary manifold and the secondary manifold distributes fuel at various pressures to both the primary and secondary manifolds. | 08-08-2013 |
20130199206 | SPEED SENSOR PROBE LOCATION IN GAS TURBINE ENGINE - A gas turbine engine includes a fan, a fan drive gear system coupled to drive the fan about an engine central axis, a compressor section including a first compressor and a second compressor and a turbine section. The turbine section includes a first turbine coupled to drive a first spool. The first spool is coupled at a first axial position to a compressor hub that is coupled to drive the first compressor. The first spool is also coupled at a second, different axial position to a fan drive input shaft that is coupled to drive the fan drive gear system. The turbine section also includes a second turbine coupled through a second spool to drive the second compressor. A sensor probe is operable to determine a rotational speed of the first spool. The sensor probe is located at a third axial position that is axially forward of the first axial position and axially aft of the second axial position. | 08-08-2013 |
20130201781 | Acoustic Acceleration of Fluid Mixing in Porous Materials - Apparatus and methods are disclosed for uniformly mixing fluid phases entrained in a porous material. A mixer may have a vessel and at least one porous material held by the vessel. At least one actuator may be acoustically coupled with at least one wall of the vessel for generating a wave. The wave effects mixing of at least two fluids in the porous material. The actuator may be a linear motor actuated with a control signal of predetermined frequency. The actuator may have a number of actuator pairs each including respective first and second actuators at respective first and second sides of the vessel. The actuators may be hinged for reciprocal movement. The actuators may be actuated to form a compression expansion wave to effect fluid motion in the porous material. | 08-08-2013 |
20130202406 | GAS TURBINE ENGINE THERMAL MANAGEMENT SYSTEM - A thermal management system for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a heat exchanger and a valve that controls an amount of a first fluid that is communicated through the heat exchanger A first sensor senses a first characteristic of a second fluid that is communicated through the heat exchanger to exchange heat with the first fluid and a second sensor senses a second characteristic of the second fluid. A positioning of the valve is based on at least one of the first characteristic and the second characteristic. | 08-08-2013 |
20130202415 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. The engine includes a combination of quantities providing beneficial operation. | 08-08-2013 |
20140112760 | REDUCTION OF EQUALLY SPACED TURBINE NOZZLE VANE EXCITATION - A reduction in excitation amplitudes affecting turbine blade durability in a turbine nozzle assembly having a plurality of vanes and turbine blades, includes: identifying a turbine blade design of the turbine nozzle assembly; performing a modal model analysis of at least one of the turbine blades in the turbine blade design; reducing aerodynamic impact by ensuring that each of the turbine blades is free of aero-excitation from an upstream flow at the vanes in an operating speed range; identifying blade natural frequencies with respect to the nozzle vanes; and modifying a trailing edge of at least one of the vanes to reduce the excitation amplitudes. | 04-24-2014 |