| UNITED TECHNOLOGIES CORPORATION Patent applications |
| Patent application number | Title | Published |
|---|
| 20120132447 | TORQUE PROTECTION DEVICE FOR FIRE BLANKET AND ASSOCIATED METHOD - A fire blanket assembly includes a thermal insulation material defining an opening, a metallic foil layer positioned along a first surface of the thermal insulation material and over at least a portion of the opening, an insert disc positioned adjacent to a second surface of the thermal insulation material opposite the first surface and further being positioned at the opening in the thermal insulation material, and a detail. The detail includes a base positioned at least partially within the opening in the thermal insulation material, a sidewall extending from the base, and a flange extending from the sidewall at a location spaced from the base, a perimeter of the flange turned at least partially away from the metallic foil layer. | 05-31-2012 |
| 20120132138 | DIMENSIONALLY STABLE DURABLE THERMAL SPRAY MASKING SYSTEM - A masking system protects portions of a part, such as a turbine engine component, to be coated. The masking system has a base, a conduit positioned on said base, a part to be coated being positioned over the conduit, and an annular plate positioned over the conduit and resting on a first portion of the part. | 05-31-2012 |
| 20120131929 | FLEXIBLE SEAL FOR GAS TURBINE ENGINE SYSTEM - A system includes a gas turbine engine, a nozzle structure located adjacent to the gas turbine engine, a flexible annular seal linking the gas turbine engine and the nozzle structure, the flexible annular seal having a plurality of corrugate sections, and a reinforcement member positioned between two of the plurality of corrugate sections. | 05-31-2012 |
| 20120131928 | FLEXIBLE SEAL FOR GAS TURBINE ENGINE SYSTEM - A system includes a gas turbine engine, a nozzle structure located adjacent to the gas turbine engine, a flexible annular seal linking the gas turbine engine and the nozzle structure, the flexible annular seal having a plurality of corrugate sections arranged about a centerline, and a connection device secured to the gas turbine engine and to the nozzle structure. The connection device is adjustable to select a length of the flexible annular seal measured along the centerline. | 05-31-2012 |
| 20120125011 | METHOD FOR USE WITH ANNULAR GAS TURBINE ENGINE COMPONENT - A method is provided for use with an annular gas turbine engine component that includes a main body and an annular seal engaged to the main body with an interference fit. The method includes rotating the annular gas turbine engine component, directing thermal energy from a heat source at the annular seal to thermally expand the annular seal relative to the main body while the annular gas turbine engine component rotates, removing the heated annular seal from the main body, providing a newly manufactured replacement detail having substantially the same configuration as the main body, and engaging the removed annular seal to the replacement detail with an interference fit. | 05-24-2012 |
| 20120112415 | ROTATING SEAL RING WITH TARGETED SPLIT SURFACE ORIENTATION - A seal assembly includes a first component, a second component that defines an outer surface and is located radially inward from the first component, a groove defined in the second component and arranged to face the first component, and a seal ring positioned between the first and second components and extending at least partially into the groove. The seal ring defines an outer diameter surface and a first lateral surface adjoining the outer diameter surface. The seal ring is split to define a first free end and a second free end, and the first and second free ends are configured to overlap along a split surface that extends from the first lateral surface to the outer diameter surface. | 05-10-2012 |
| 20120112394 | DAMPENING DEVICE - An assembly for dampening vibration movement between a cylinder and a concentric duct, where the cylinder is positioned radially inside the duct, includes a damper attached to the duct. The damper includes a housing attached to the duct, a contact member positioned at least partially inside the housing and having a contact surface adjacent an outer surface of the cylinder, and a cap threadedly engaged with the housing for applying a load on the contact member in a direction toward the cylinder. | 05-10-2012 |
| 20120110975 | IGNITER WITH INTEGRAL PRESSURE SENSING LINE - An igniter device for a combustor system includes an igniter housing and a pressure sense passage. The igniter housing surrounds an electrode and an insulating body. The igniter has a tip to be positioned within a chamber for combustion. The pressure sense passage is attached to an exterior surface of the igniter housing. The pressure sense passage is configured to direct fluid to a pressure sensor. | 05-10-2012 |
| 20120110848 | LOW AND EXTRA LOW SULFUR ALLOYS FOR REPAIR - A method for repairing, refurbishing, or replacing a turbine engine component or sub-component includes the steps of providing a turbine engine component or sub-component having a site to be repaired, refurbished, or replaced providing a repair or replacement material having a sulfur content, which sulfur content is less than 10 ppm, and applying the repair or replacement material to the site on the turbine engine component to effect the repair, the refurbishment, or the replacement. | 05-10-2012 |
| 20120107098 | GAS TURBINE ENGINE ROTOR TIE SHAFT ARRANGEMENT - Independent axial preloading of compressor and turbine disks in a gas turbine engine rotor is achieved by a dual tie shaft arrangement wherein a forward tie shaft engages forward and aft portions of a compressor hub within which the compressor disks are disposed to axially and compressively preload the compressor disks between the forward and aft portions of the compressor hub. An independent tie shaft engages an aft portion of a turbine hub within which turbine disks are disposed and an aft portion of the compressor hub to axially and compressively preload the turbine disks between the turbine and compressor hubs. | 05-03-2012 |
| 20120107095 | GAS TURBINE ENGINE TRIM BALANCE - A trim balance for a gas turbine engine having a hollow rotatable rotor shaft includes an elongate support rod adapted for longitudinal accommodation within and mounting on an interior surface of the rotor shaft and one or more apertured balance weights supported on the support rod which extends through the weight apertures and a weight retainer for retaining the weights on the support rod. | 05-03-2012 |
| 20120107087 | METHOD FOR VARYING THE GEOMETRY OF A MID-TURBINE FRAME - A method for varying the geometry of a mid-turbine frame includes detecting a strain in a mid-turbine frame with a piezoelectric material; applying a deformation voltage to the piezoelectric material as a function of the detected strain; deforming the piezoelectric material to actuate an actuation plate; and repositioning an engine casing through the actuation of the actuation plate. | 05-03-2012 |
| 20120102960 | GAS TURBINE ENGINE SYSTEMS INVOLVING COOLING OF COMBUSTION SECTION LINERS - Gas turbine engine systems involving cooling of combustion section liners are provided. A representative liner includes: an outer side, an inner side, an upstream end, and a downstream end, the outer side being configured to face away from a combustion reaction, the inner side being configured to face the combustion reaction; a cooling air channel, a portion of the cooling air channel being located proximate the downstream end; and cooling holes formed through the inner side of the liner, the cooling holes being in fluid communication with the cooling air channel such that cooling air provided to the cooling air channel is directed through the cooling holes and to the inner side of the liner such that at least a portion of the inner side of the liner receives cooling air despite a corresponding portion located on the outer side of the liner being obstructed from receiving cooling air. | 05-03-2012 |
| 20120100299 | THERMAL SPRAY COATING PROCESS FOR COMPRESSOR SHAFTS - A process for forming a dense abrading thermally insulating coating on a rotor shaft in a gas turbine engine is described. The process comprises fixturing the rotor shaft to allow it to rotate about its axis and plasma spraying the coating on the rotor shaft. The coating comprises a zirconia based ceramic top coat layer on a metallic bond coat. | 04-26-2012 |
| 20120099992 | ABRASIVE ROTOR COATING FOR FORMING A SEAL IN A GAS TURBINE ENGINE - A compressor rotor seal includes a bare radial inward end of a rotor vane in proximity to a rotor surface coated with an abrasive material. | 04-26-2012 |
| 20120099985 | ABRASIVE CUTTER FORMED BY THERMAL SPRAY AND POST TREATMENT - A seal for a gas turbine engine component having an airfoil with a radial outward end and a radial inward end, one of which is bare metal. A seal member is adjacent to the bare metal end of the airfoil wherein the seal member is coated with a abrasive layer having a laser engraved surface with the top of the surface has less than about 5% of the surface area of the base of the abrasive layer. | 04-26-2012 |
| 20120099973 | LOW DENSITY ABRADABLE COATING WITH FINE POROSITY - A low density abradable coating for use as a seal material is formed by depositing a metal bond coat followed by depositing a low density, porous, abradable metal seal layer on the bond coat. The seal layer is formed by co-depositing metallic sponge particles and metal precursor particles. The metal precursor particles decompose during heat treatment leaving fine pores distributed throughout the microstructure. | 04-26-2012 |
| 20120099972 | ROUGH DENSE CERAMIC SEALING SURFACE IN TURBOMACHINES - An abrasive coating is formed on a rotor shaft that rotates with respect to cantilevered vanes. The coating is formed by thermal spray techniques and comprises a ceramic layer on a metal bond coat. The coating surface is roughened by crush grinding or by grit blasting to increase the abradability of the surface. | 04-26-2012 |
| 20120099971 | SELF DRESSING, MILDLY ABRASIVE COATING FOR CLEARANCE CONTROL - An abrasive coating for rotor shafts that interact with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating including a metal bond coat layer on the rotor shaft, and an abrasive top coating bond coat layer for contact with vanes during operation of the rotor shaft, the abrasive coating including a plurality of abrasive grit particles in a matrix. the abrasive grit particles are selected from the group consisting of cubic boron nitride (CBN), zirconia, alumina, silicon carbide, diamond and mixtures thereof. | 04-26-2012 |
| 20120099970 | FRIABLE CERAMIC ROTOR SHAFT ABRASIVE COATING - An abrasive coating on a rotor shaft interacts with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating includes a metal bond coat, and an abrasive layer containing a plurality of abrasive CBN grit particles in a ceramic matrix. | 04-26-2012 |
| 20120099968 | ABRASIVE ROTOR SHAFT CERAMIC COATING - An abrasive coating on a rotor shaft interacts with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating includes a metal bond coat, and an abrasive layer containing a plurality of abrasive grit particles in a ceramic matrix. The grit particles may be selected from cubic boron nitride (CBN), zirconia, alumina, silicon carbide, diamond and mixtures thereof. | 04-26-2012 |
| 20120097295 | HALOGEN-FREE TRIVALENT CHROMIUM CONVERSION COATING - Trivalent chromium conversion coatings are provided on a metal substrate wherein the trivalent chromium conversion coating has a halogen content of 1 atom % maximum. | 04-26-2012 |
| 20120095590 | REVERT BLEND ALGORITHM AND APPARATUS USING THE ALGORITHM - A revert alloy is used in the manufacture of a target alloy having a different composition. The weight percent of the elemental constituents of this first alloy and a second or target alloy are obtained, and the ratio of the percentage by weight of each elemental constituent in the first alloy to the second alloy is determined. The lowest ratio is used to determine the amount of the first alloy that is melted and the necessary elemental constituents that are added to the melt to produce the desired composition of the target alloy. The melt is solidified to produce the target alloy. | 04-19-2012 |
| 20120090329 | SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES - Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative spool assembly for a gas turbine engine, which incorporates a compressor, a turbine and a gear assembly, includes: a shaft operative to be driven by the turbine; a first spool segment operative to couple the shaft to the compressor; and a second spool segment operative to couple the shaft to the gear assembly. The first spool segment and the second spool segment are not coupled to each other. | 04-19-2012 |
| 20120085528 | INTERDEPENDENT LUBRICATION SYSTEMS - A heat exchange system for use in operating equipment having a plurality of subsystems in each of which one of a plurality of working fluids is utilized to provide selected operations with there being an air and working fluid heat exchanger providing controlled cooling to cool at least one of the plurality of working fluids in its corresponding subsystem. In addition, a coupling heat exchanger is also provided connected to two of the subsystems to pass there working fluids therethrough, including the subsystem with the air and working fluid heat exchanger, to allow one of the connected subsystems to aid in cooling the other. | 04-12-2012 |
| 20120080272 | REDUCED GULP FLUID RESERVOIR - An assembly includes a reservoir for holding a fluid, and a scavenge passage connected to the reservoir at a reservoir inlet. The scavenge passage returns the fluid from a delivery location to the reservoir. A bleed passage is connected in fluid communication between the reservoir and the scavenge passage. | 04-05-2012 |
| 20120070276 | TURBINE SHROUD THERMAL DISTORTION CONTROL - A shroud suitable for use in combination with an adjacent rotor blade includes a leading portion comprising a first thickness, and a trailing portion adjacent to the leading portion and comprising a second thickness, wherein the second thickness is less than the first thickness so that stiffness in the leading portion is greater than in the trailing portion and radial expansion by thermal growth of the leading portion of the shroud is constrained. | 03-22-2012 |
| 20120068003 | WIRELESS SENSOR FOR AN AIRCRAFT PROPULSION SYSTEM - A method is provided for wirelessly monitoring an aircraft propulsion system having a plurality of rotor blades, at least some of which rotor blades have an internal compartment. The method includes: providing a primary transducer having a primary coil, and one or more secondary transducers, each secondary transducer having a secondary coil connected to a sensor, wherein the secondary transducers are respectively disposed within the internal compartments; wirelessly powering the secondary transducers using a magnetic field generated by the primary transducer; monitoring at least one operational parameter from within each internal compartment using a respective sensor; and transmitting output signals from the secondary transducers to the primary transducer through the magnetic field, wherein the output signals are indicative of the monitored parameters. | 03-22-2012 |
| 20120067049 | SYSTEMS AND METHODS FOR POWER GENERATION FROM MULTIPLE HEAT SOURCES USING CUSTOMIZED WORKING FLUIDS - A power generating system in one embodiment employs a Rankine Cycle system that is coupled to multiple heat sources. The Rankine cycle system includes a customized working fluid that comprises a mixture of a plurality of constituent fluids, the selection of which causes the mixture to exhibit a working fluid profile. In one embodiment, the working fluid profile includes a temperature glide portion selected and optimized based on operating conditions of the heat sources, wherein the temperature glide portion includes a constituent phase point at which one of the constituent fluids undergoes a phase change before the other constituent fluids of the mixture. | 03-22-2012 |
| 20120063888 | ABRADABLE COATING WITH SAFETY FUSE - An abradable coating for interaction with tips of airfoils (such as vanes or blades of a compressor of a gas turbine engine) includes a metal bond coat thermal, a ceramic layer, and an abradable layer. The ceramic layer on the metal bond coat provides insulation and acts as a fuse that is adapted to spall off upon high rub interaction. The abradable coating on the ceramic layer contacts the tips of the airfoils during operation of the compressor. The abradable coating is sufficiently abradable to roundup the coating by contact with the airfoil tips. | 03-15-2012 |
| 20120060505 | ADAPTIVE CONTROL FOR A GAS TURBINE ENGINE - A method for controlling a gas turbine engine includes: generating model parameter data as a function of prediction error data, which model parameter data includes at least one model parameter that accounts for off-nominal operation of the engine; at least partially compensating an on-board model for the prediction error data using the model parameter data; generating model term data using the on-board model, wherein the on-board model includes at least one model term that accounts for the off-nominal operation of the engine; respectively updating one or more model parameters and one or more model terms of a model-based control algorithm with the model parameter data and model term data; and generating one or more effector signals using the model-based control algorithm. | 03-15-2012 |
| 20120060466 | AIR-OIL HEAT EXCHANGER - A heat exchange system for use in fluid operated equipment to provide air and working fluid heat exchanges to cool the working fluid in airstreams on a stream side of a wall. An actuator is mounted to be substantially located on a side of the wall opposite the stream side thereof having a positionable motion effector. A heat exchanger core having a plurality of passageway structures therein to enable providing the working fluid to, and removal therefrom. The heat exchanger core is mounted on the motion effector so as to be extendable and retractable thereby through the opening for selected distances into that region to be occupied by the airstreams. | 03-15-2012 |
| 20120060370 | SYSTEM AND METHOD FOR REPAIRING A GAS TURBINE ENGINE COMPONENT - A method for repairing a defect in a component having a first major surface and a second major surface, a thickness defined between the major surfaces, and an edge adjoining both the major surfaces is characterized by forming a slot cut entirely through the thickness of the component between the first and second major surfaces, the slot extending inward from the edge of the component towards an interior of the component to a depth from the edge sufficient to remove at least a substantial part of the defect, the slot defining an open end at the edge of the component, a closed end within the component, a width, a radius formed by the closed end, and opposing chamfered edges at the open end. | 03-15-2012 |
| 20120057982 | TURBINE VANE AIRFOIL - An airfoil comprises an external surface formed in substantial conformance with a plurality of cross sections. Each of the cross sections has an axial chord and an offset at a corresponding fractional span as set forth in Table 1. Each of the cross sections is described by a set of Cartesian coordinates defined at the corresponding fractional span as set forth in Table 3. | 03-08-2012 |
| 20120057981 | TURBINE BLADE AIRFOIL - An airfoil comprises an external surface formed in substantial conformance with a plurality of cross sections. Each of the cross sections has an axial chord and an offset at a corresponding fractional span as set forth in Table 2. Each of the cross sections is described by a set of Cartesian coordinates defined at the corresponding fractional span as set forth in Table 4. | 03-08-2012 |
| 20120055647 | Airfoil Casting Methods - In the castings of a turbine element, a sheet casting core is assembled to a feedcore. The sheet casting core and feedcore are placed in a die. A sacrificial pattern material is molded over the casting core and feedcore to form a pattern including an airfoil. The sheet casting core extends from at or adjacent a trailing edge of the airfoil. The sheet casting core has a first array of open areas and a second array of portions interspersed with the open areas. A first portion of the die has a third array of projections contacting the second array. | 03-08-2012 |
| 20120045319 | IMPINGEMENT COOLING FOR TURBOFAN EXHAUST ASSEMBLY - A system comprises a bypass duct, a turbine case, a turbine exhaust ring and a splash plate. The turbine case is formed along a radially inner margin of the bypass duct, and the turbine exhaust ring is coaxially disposed within the turbine case. The splash plate extends axially along the turbine exhaust ring, radially spaced between the turbine exhaust ring and the turbine case. There are cooling fluid apertures in the turbine case to provide cooling fluid flow from the bypass duct onto the splash plate, and impingement holes in the splash plate to provide impingement flow onto the turbine exhaust ring. | 02-23-2012 |
| 20120039707 | METHOD OF REPAIRING KNIFE EDGE SEALS - A knife edge seal assembly includes at least one knife edge and at least one honeycomb ring seal. The at least one knife edge has material removed to a first knife edge radial height and the least one honeycomb ring seal has material removed to a finished radial thickness for sealable engagement with the at least one knife edge. The finished radial thickness of the replacement honeycomb ring seal establishes a first radial gap between an inner annular surface of the replacement honeycomb ring seal and the knife edge before the knife edge seal assembly begins rotating in the gas turbine. | 02-16-2012 |
| 20120034491 | CMAS RESISTANT TBC COATING - A process for forming a coating system on a turbine engine component comprises the steps of providing a substrate, depositing a thermal barrier coating on the substrate, depositing a reactive with known CMAS reaction kinetics on the thermal barrier coating, and activating the reactive layer prior to the component being placed in service. As a result of the foregoing process, there is provided a turbine engine component which has a substrate, a thermal barrier coating deposited on the substrate, a reactive layer deposited on the thermal barrier coating, which reactive layer has known CMAS reaction kinetics and is activated prior to the turbine engine component entering into service. | 02-09-2012 |
| 20120034100 | RECESSED METERING STANDOFFS FOR AIRFOIL BAFFLE - An internally cooled airfoil comprises an airfoil body, a baffle and a plurality of standoffs. The airfoil body is shaped to form leading and trailing edges, and pressure and suction sides surrounding an internal cooling channel. The baffle is disposed within the internal cooling channel and comprises a liner body having a perimeter shaped to correspond to the shape of the internal cooling channel and to form a cooling air supply duct. The baffle includes a plurality of cooling holes extending through the liner body to direct cooling air from the supply duct into the internal cooling channel. The standoffs maintain minimum spacing between the liner body and the airfoil body. The standoffs are recessed into a surface of either the baffle or the airfoil body such that a height of the standoffs is greater than the spacing. | 02-09-2012 |
| 20120027601 | ROTATABLE COMPONENT MOUNT FOR A GAS TURBINE ENGINE - Radial shifting of a rotatable component in a gas turbine engine is prevented by radially offsetting overlying mounting apertures in said component and a base or mounting flange therefor such that fasteners received within said overlying apertures are radially interference fit within the apertures thereby eliminating the necessity of machining or otherwise forming the apertures to an exact fit with the fasteners. | 02-02-2012 |
| 20120027574 | BLADE OUTER AIR SEAL AND REPAIR METHOD - An article of manufacture has a body formed in part of a first metal alloy and in part of a second metal alloy, the second metal alloy having a thermal coefficient of expansion that is less than the thermal coefficient of expansion of the first metal alloy. A BOAS segment for a gas turbine engine is disclosed wherein the formation of cracks due to thermal mechanical fatigue in the body of the disclosed BOAS segment is minimized, if not eliminated, through a unique construction of the disclosed BOAS segment, whether original equipment manufacture or a repaired blade outer air seal. A method for manufacture of a BOAS segment and a method for modifying a BOAS segment are disclosed. | 02-02-2012 |
| 20120027570 | SHAFT ASSEMBLY FOR A GAS TURBINE ENGINE - A shaft assembly for a gas turbine engine includes a shaft, an annular fluid collection cavity, and a bearing. The shaft is rotatable around an axially extending shaft centerline. The annular fluid collection cavity is formed by an axial panel extending between a forward radial panel and an aft radial panel. The aft radial panel includes a plurality of fluid apertures, each fluid aperture having a center disposed at a first radius from the shaft centerline. The bearing includes an inner race having a shaft mounting surface and a plurality of axially extending race fluid passages. The shaft mounting surface is engaged with an exterior surface of the shaft, and is disposed at a second radius from the shaft centerline, wherein the second radius is less than the first radius. At least some of the axially extending race fluid passages are fluidly connected with at least some of the fluid apertures in the aft radial panel. | 02-02-2012 |
| 20120020797 | VORTEX REDUCING DEVICE FOR A GAS TURBINE ENGINE - A gas turbine engine has a vortex reducing device therein that includes a retainer and paddles. The retainer is arcuate in shape and has a plurality of circumferentially spaced slots that extend through it. The paddles are circumferentially spaced about the retainer and extend outward thereform. Each paddle is disposed between adjacent slots. | 01-26-2012 |
| 20120015109 | FLUOROPOLYMER-CONTAINING FILMS FOR USE WITH POSITIVE-DISPLACEMENT FLUID PUMPS - A positive-displacement pump comprising a housing, a rotatable component disposed at least partially within the housing, and a film disposed at least partially between the housing and the rotatable component, where the film comprises a reinforcing material and a fluoropolymer material. | 01-19-2012 |
| 20120011824 | INTEGRAL LUBRICATION TUBE AND NOZZLE COMBINATION - An assembly includes a tube fitting integrally joined with a nozzle and adapted with a plurality of passages to spray lubricating oil at components within a bearing compartment of a gas turbine engine. In another aspect, the tube fitting and the nozzle are attached to a lubrication tube such that the tube fitting and nozzle can be inserted or removed from the bearing compartment along with the lubrication tube without disassembling the bearing compartment. | 01-19-2012 |
| 20120007286 | LOW TRANSIENT AND STEADY STATE THERMAL STRESS DISK SHAPED COMPONENTS - A process for manufacturing a disk shaped component comprising fabricating a disk shaped component using a composite material having at least a first material and a second material, wherein the first material is disposed at and proximate to a center portion of the disk shaped component and the second material is disposed at and proximate to a rim of the disk shaped component, wherein the first material comprises a first coefficient of thermal expansion, a first stress value and a first oxidation resistance, and the second material comprises a second coefficient of thermal expansion, a second stress value and a second oxidation resistance, wherein the first coefficient of thermal expansion is greater than the second coefficient of thermal expansion, the first stress value is greater than the second stress value and the first oxidation resistance is less than the second oxidation resistance. | 01-12-2012 |
| 20120006034 | HYBRID COOLING OF A GAS TURBINE ENGINE - A method for cooling components of a gas turbine engine, having first and second sets of components that require cooling, includes removing thermal energy from cooling air, directing the cooling air from which thermal energy has been removed to the first set of components of the gas turbine engine, and transferring thermal energy away from the second set of components of the gas turbine engine using a vaporization cycle of a vapor cooling assembly. Use of the vaporization cycle offsets thermal energy removed from the cooling air. | 01-12-2012 |
| 20120000967 | LAYERED STRUCTURES WITH INTEGRAL BRAZING MATERIALS - A layered structure comprising a base structure having a major surface, and a brazing layer secured to the major surface of the base structure, where the brazing layer is applied to the major surface prior to positioning the layered structure in contact with a turbine engine component. | 01-05-2012 |
| 20110315337 | Investment Casting Cores and Methods - A method involves forming a core assembly. The forming includes deforming a wire, the deforming including increasing a transverse linear dimension along at least one portion of the wire. The wire is assembled to a ceramic core | 12-29-2011 |
| 20110315336 | Contoured Metallic Casting Core - A method for manufacturing an investment casting core uses a metallic blank having a thickness between parallel first and second faces less than a width and length transverse thereto. The blank is locally thinned from at least one of the first and second faces. The local thinning forms a taper on a leading portion of the RMC. The blank is through-cut across the thickness. The blank is inserted into the leading portion into a slot in a pre-formed ceramic core. | 12-29-2011 |
| 20110314824 | SWIRLER, FUEL AND AIR ASSEMBLY AND COMBUSTOR - An air swirler, a fuel and air admission assembly, and a staged combustor are disclosed. The staged combustor may be equipped with the fuel and air admission assemblies incorporating the air swirlers for use in gas turbine engines, such as for example gas turbine engines powering aircraft having supersonic cruise capability. | 12-29-2011 |
| 20110314823 | GAS TURBINE COMBUSTOR LINER CAP ASSEMBLY - A combustor liner cap assembly is provided for use in a multiple fuel nozzle combustor of a gas turbine. The combustor liner cap assembly includes a tube plate having a plurality of fuel nozzle openings and a plurality of open ended premix tubes extending aft from the tube plate. Each premix tube has a forward end having a forward edge, the forward end being received in a corresponding one of the fuel nozzle openings. Each premix tube is secured to the tube plate aft of the forward edge of the forward end of the premix tube, for example, metallurgically bonded, such as by welding or brazing, to the tube plate or threaded into the tube plate. | 12-29-2011 |
| 20110314818 | CASCADED CONDENSER FOR MULTI-UNIT GEOTHERMAL ORC - A pair of organic rankine cycle systems are connected in series with the geothermal fluid passing first through an evaporator of the first system and then through an evaporator of the second system before returning to a sink. Similarly, the cooling tower is arranged to provide cooling water to pass first through the condenser in one system and then through the condenser of the other system, to reduce the total flow required and the size of associated cooling hardware. | 12-29-2011 |
| 20110309066 | ENGINE INLET ICE PROTECTION SYSTEM HAVING EMBEDDED VARIABLE WATT DENSITY HEATERS - A system includes a first heater located at the leading edge of a gas turbine structural member, a second heater located aft of the first heater, and a third heater located aft of the second heater. The first, second and third heaters are electrically-powered to prevent icing of the gas turbine structural member. Each of the heaters has a Watt density, and the Watt densities of the heaters differ from one another as a function of a magnitude of a cooling coefficient for airflow passing the vicinity of each heater. | 12-22-2011 |
| 20110308290 | METHOD FOR ULTRASONIC PEENING OF GAS TURBINE ENGINE COMPONENTS WITHOUT ENGINE DISASSEMBLY - A method for peening a component of a gas turbine engine, or a module of a gas turbine engine, without removing the component from the assembled gas turbine engine or module includes blending the damaged area of the component, determining parameters for peening the damaged component, and peening the damaged component without removing it from the engine. | 12-22-2011 |
| 20110308254 | Turbine Engine Fuel Injector - A gas turbine engine is piloted with a pilot flow of fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A second additional flow of the fuel is vaporized and delivered to the combustor as a vapor. A fuel injector may have passageways associated with each of the three flows. | 12-22-2011 |
| 20110305892 | POROUS PROTECTIVE COATING FOR TURBINE ENGINE COMPONENTS - A method for coating a substrate of a turbine engine component, the method comprising cold spray depositing a metal-based material onto a surface of the substrate, and heating the deposited metal-based material to increase the porosity of the deposited metal-based material. | 12-15-2011 |
| 20110305576 | GAS TURBINE ENGINE BLADE MOUNTING ARRANGEMENT - A mounting arrangement for a gas turbine engine blade having a root employs a lightweight, simple and economical leaf spring spacer which biases the blade root in a radially outward direction to minimize unwanted movement of the root within a conforming slot in a blade hub under conditions such as windmilling when centrifugal force alone is inadequate to tightly seal the root within the slot. | 12-15-2011 |
| 20110300342 | Ceramic Coating Systems and Methods - An article has a metallic substrate. The substrate has a first surface region and a plurality of blind recesses along the first surface region. The substrate has perimeter lips at the openings of the plurality of recesses and extending partially over the respective associated recesses. A ceramic coating is along the first surface region. | 12-08-2011 |
| 20110293826 | Apparatus and Method for Coating Internal Surfaces of a Turbine Engine Component - A coating apparatus for coating interior surfaces of internal passages of a workpiece, such as a turbine engine component is provided. The coating apparatus comprises a vessel defining an internal space, a plate dividing the internal space into a first chamber and a second chamber, a divider supporting the plate, and the plate having at least one cut-out for receiving a portion of the workpiece having the internal passages to be coated. | 12-01-2011 |
| 20110287188 | PROCESSES FOR APPLYING A CONVERSION COATING WITH CONDUCTIVE ADDITIVE(S) AND THE RESULTANT COATED ARTICLES - A process for coating an article includes the steps of contacting an article with a first solution to produce a coated article, the first solution includes a solvent and at least one non-conductive material comprising at least one oxide of a metal; contacting with a second solution the coated article having at least one surface with a non-conductive material layer, the second solution includes a solvent and at least one conductive material comprising at least one of the foregoing: graphite, metals, conductive ceramics, semi-conductive ceramics, intermetallic compounds, and mixtures thereof; and drying the coated article having at least one surface with a non-conductive material layer having the at least one conductive material in contact with at least one surface of the non-conductive material layer and the at least one surface of the article. | 11-24-2011 |
| 20110281107 | LAYERED THERMAL BARRIER COATING WITH BLENDED TRANSITION AND METHOD OF APPLICATION - A multilayer coating includes a bond coat layer and a first barrier layer applied on the bond coat 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. | 11-17-2011 |
| 20110274559 | Turbine Airfoil with Body Microcircuits Terminating in Platform - A turbine engine component includes a platform and one or more microcircuit cooling passages embedded within one or more walls of an airfoil portion of the component. Each microcircuit cooling passage terminates within the thickness of the platform so as to provide cooling to the initial 10% span of the airfoil portion. Each microcircuit cooling passage has an inlet for receiving cooling fluid, which inlet is also embedded within the platform. | 11-10-2011 |
| 20110268578 | HIGH PITCH-TO-CHORD TURBINE AIRFOILS - A low pressure turbine engine component for use in an engine, for propelling a vehicle such as an aircraft is formed by a disk portion and a plurality of low pressure turbine blades extending outwardly from the disk portion. Each of the low pressure turbine blades has an airfoil portion with an axial chord length and a trailing edge. The low pressure turbine blades are spaced apart so that there is a pitch-to-chord ratio greater than 1.4, wherein the pitch is a distance between the trailing edges of adjacent ones of the low pressure turbine blades and the chord is the axial chord length of the blades. | 11-03-2011 |
| 20110267451 | ON-THE-FLY DIMENSIONAL IMAGING INSPECTION - A method and system are provided for inspecting a plurality of target features arrayed in spaced arrangement on a surface of a target object, such as but not limited to inspection of the location of cooling air holes in the surface of a turbine blade or vane. | 11-03-2011 |
| 20110262274 | ENGINE ASSEMBLED SEAL - A system for creating a seal between a damper and a platform of a turbine engine component includes a turbine engine component having an airfoil portion, a platform, and a fir tree for joining the turbine engine component to a rotor, a damper located in an area beneath the platform, a seal having a sealing surface which seats against an underside of the platform, which seal has a seal retention feature which bends into contact with an underside of the damper, and which seal with the seal retention feature has a center of gravity which allows the seal retention feature to bend up as result of rotational movement of the rotor. A method for creating the seal is also described. | 10-27-2011 |
| 20110256421 | METALLIC COATING FOR SINGLE CRYSTAL ALLOYS - A metallic coating for use in a high temperature application is created from a nickel base alloy containing from 5.0 to 10.5 wt % aluminum, from 4.0 to 15 wt % chromium, from 2.0 to 8.0 wt % tungsten, from 3.0 to 10 wt % tantalum, and the balance nickel. The metallic coating has particular utility in protecting single crystal superalloys used in high temperature applications such as turbine engine components. | 10-20-2011 |
| 20110248451 | WEAR-RESISTANT COATING - A seal assembly for a gas turbine engine includes a first seal member having a first surface, a second seal member having a second surface, with the second surface configured to generally abut at least a part of the first surface. At least a portion of at least one of the first surface and the second surface includes a coating that includes about 30 to about 80 weight percent of a hard carbide material, and about 20 to about 70 weight percent of lubricating material incorporated with the hard carbide material. The coating defines overlapping lenticular particles. | 10-13-2011 |
| 20110247728 | HALOGEN-FREE TRIVALENT CHROMIUM CONVERSION COATING - Trivalent chromium conversion coatings are provided on a metal substrate wherein the trivalent chromium conversion coating has a halogen content of 1 atom % maximum. | 10-13-2011 |
| 20110247343 | GAS TURBINE ENGINE SYSTEMS AND RELATED METHODS INVOLVING VANE-BLADE COUNT RATIOS GREATER THAN UNITY - Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes. | 10-13-2011 |
| 20110243719 | NON-CIRCULAR AFT NACELLE COWLING GEOMETRY - A nacelle for a turbofan propulsion system that extends along a centerline includes a forward cowling and an aft cowling. To improve the fit of a turbofan propulsion system in the space between the wing and ground of a fixed-wing aircraft, the aft cowling of the nacelle is modified. The aft cowling has a non-circular cross-sectional geometry disposed in a plane substantially perpendicular to the centerline. The non-circular cross-sectional geometry includes a radially recessed section disposed between first and second curved sections. The first and the second curved sections each have a radius that is greater than a radial distance between the centerline and a center point of the radially recessed section. | 10-06-2011 |
| 20110241295 | METHOD OF FORMING A SEAL ELEMENT - A method is provided for forming a seal element on a substrate surface. A layer of base material is formed on the substrate surface and a plurality of crenellations are formed in the layer of base material prior to forming a layer of seal material on the crenellated layer of base material. The method may be applied to making an original seal element or to repairing an existing seal element. The resulting seal includes a layer of base material bonded to a substrate and having a plurality of crenellations formed in a face surface of the layer of a base material and a layer of a seal material bonded thereto, the seal material filling the crenellations. | 10-06-2011 |
| 20110239934 | MASKING APPARATUS - A masking apparatus is provided for masking a workpiece, for example a turbine blade, to protect portions thereof from the deposition of coating material thereon during selective application of a coating material to other portions of the workpiece. The masking apparatus includes an enclosure formed of a first mask half body defining a first cavity half and a second mask half body defining a second cavity half. The first mask half body and the second mask half body mate in assembly to form a workpiece cavity it which the workpiece may be enclosed. | 10-06-2011 |
| 20110239660 | MOUNTING ARRANGEMENT FOR GAS TURBINE ENGINE ACCESSORIES AND GEARBOX THEREFOR - Gas turbine engine accessories are mounted on an accessory gearbox therefor which is in turn mounted on a case of the engine such that the accessories extend from the gearbox in directions generally parallel to circumferential tangents to the engine case for compactness, ease in installation and removal of the accessories, reduction of the weight in lines extending from the accessories and minimization of the risk of damage to such accessories and gearbox from the fan blades which may separate from a fan hub upon encountering birds, ice, or other foreign objects. | 10-06-2011 |
| 20110236188 | BLADE OUTER SEAL FOR A GAS TURBINE ENGINE - A blade outer air seal for a gas turbine engine is provided. The blade outer air seal includes a body having an outer radial surface, an inner radial surface, and a plurality of cooling air apertures. The body extends between a forward edge and an aft edge. The inner radial surface includes at least one first seal section, at least one second seal section, and a riser extending radially between the first seal section and the second seal section. Each of the plurality of cooling air apertures extends between the outer radial surface and the riser, and each cooling air aperture has an exit configured to direct cooling air substantially parallel to the second seal section of the inner radial surface. | 09-29-2011 |
| 20110236179 | SEAL CLEARANCE CONTROL ON NON-COWLED GAS TURBINE ENGINES - A method for controlling turbine blade tip seal clearance on a non-cowled gas turbine engine includes the step of directing a flow of relatively hot, pressurized air in contact with an exterior surface of the engine core casing circumscribing the turbine. The flow of relatively hot, pressurized air may be selectively directed in contact with the exterior surface of the engine core casing circumscribing the turbine when it is desired to expand the turbine shroud radially outward, such as during transient operation of the gas turbine engine. | 09-29-2011 |
| 20110231021 | DESIGN AND CONTROL OF ENGINEERING SYSTEMS UTILIZING COMPONENT-LEVEL DYNAMIC MATHEMATICAL MODEL WITH SINGLE-INPUT SINGLE-OUTPUT ESTIMATOR - A control system comprises an actuator, a control law and a processor. The actuator positions a control surface and the control law controls the actuator. The processor comprises an open loop module, a corrector, a comparator, and an estimator, and generates model output to direct the control law. The open loop module generates the model output as a function of a model state and a model input. The corrector generates a corrector output as a function of the model output. The comparator generates an error by comparing the corrector output to the model input. The estimator generates the model state as a function of the error, such that the error is minimized as a function of single-input, single-output gain matrix. | 09-22-2011 |
| 20110230981 | DESIGN AND CONTROL OF ENGINEERING SYSTEMS UTILIZING COMPONENT-LEVEL DYNAMIC MATHEMATICAL MODEL WITH MULTIPLE-INPUT MULTIPLE-OUTPUT ESTIMATOR - A control system comprises an actuator, a control law and a processor. The actuator positions a control surface and the control law controls the actuator. The processor comprises an open loop module, a corrector, a comparator, and an estimator, and generates an output vector to direct the control law. The open loop module generates the output vector as a function of a state vector and an input vector. The corrector generates a corrector vector as a function of the output vector. The comparator generates an error vector by comparing the corrector vector to the input vector. The estimator generates the state vector as a function of the error vector, such that the error vector is minimized. | 09-22-2011 |
| 20110229334 | COMPOSITE LEADING EDGE SHEATH AND DOVETAIL ROOT UNDERCUT - An airfoil having a composite blade formed from a plurality of plies and having a leading edge and a root for attachment to an engine. The blade has a decreased number of plies at the junction of the blade leading edge and the root. A metallic sheath is attached to the leading edge, wherein the sheath has a portion proximate the junction of sufficient thickness to restore at least a portion of the decreased number of plies | 09-22-2011 |
| 20110223356 | COATING APPARATUS AND METHOD WITH INDIRECT THERMAL STABILIZATION - An apparatus includes a work piece support for holding and selectively rotating a work piece, a coating delivery apparatus configured to apply a coating material to the work piece, a susceptor positioned adjacent to the work piece support, and a first electron gun configured to direct a first electron beam at the susceptor such that the susceptor radiates heat toward the work piece. | 09-15-2011 |
| 20110223355 | THERMAL STABILIZATION OF COATING MATERIAL VAPOR STREAM - A coating system includes a work piece, a coating delivery apparatus configured to apply a coating material to the work piece in a plasma-based vapor stream, and a first electron gun configured to direct a first electron beam at the plasma-based vapor stream for adding thermal energy to the coating material in the plasma-based vapor stream. | 09-15-2011 |
| 20110223354 | HIGH PRESSURE PRE-OXIDATION FOR DEPOSITION OF THERMAL BARRIER COATING - An apparatus for coating a work piece includes a process chamber, a coating material supply apparatus located at least partially within the process chamber for delivering a coating material to the work piece, a pre-heater assembly adjoining the process chamber, and a support for holding the work piece. The pre-heater assembly includes a housing that opens to the process chamber, a susceptor comprising a ceramic material positioned at least partially within the housing, and a pre-heater electron gun configured to configured to direct an electron beam at the susceptor such that the susceptor radiates heat toward the work piece. The support is movable to selectively move the work piece between a first position within the housing of the pre-heater assembly and a second position within the process chamber and outside the housing of the pre-heater assembly. | 09-15-2011 |
| 20110223353 | HIGH PRESSURE PRE-OXIDATION FOR DEPOSITION OF THERMAL BARRIER COATING WITH HOOD - An apparatus for coating a work piece includes a process chamber, a coating material supply apparatus located at least partially within the process chamber for delivering a coating material to the work piece, a pre-heater assembly adjoining the process chamber, and a support for holding the work piece. The pre-heater assembly includes a housing that opens to the process chamber, a thermal hood positioned within the housing and configured to reflect thermal energy for reflecting thermal energy toward the work piece. The support is movable to selectively move the work piece between a first position within the housing of the pre-heater assembly and a second position within the process chamber and outside the housing of the pre-heater assembly. | 09-15-2011 |
| 20110223317 | DIRECT THERMAL STABILIZATION FOR COATING APPLICATION - A coating system includes a first work piece, a work piece support for holding the first work piece, a plasma-based coating delivery apparatus configured to apply a coating material to the first work piece in a plasma-based vapor stream, and a first electron gun configured to direct a first electron beam at the first work piece while the plasma-based coating delivery apparatus applies the coating to the first work piece for heating the first work piece being coated, wherein the first electron gun is configured to direct the first electron beam at a region of the first work piece facing away from the plasma-based coating delivery apparatus. | 09-15-2011 |
| 20110223027 | COMPOSITE FAN BLADE DOVETAIL ROOT - A composite fan blade having an airfoil with a leading edge and a trailing edge and a dovetail root includes a first pressure face and second pressure face, a lower horizontal face connecting the pressure faces at the bottom of the dovetail and a first and a second dovetail runout fillet connecting the pressure faces to the airfoil. The pressure faces are each angled outward at about 65 degrees to about 75 degrees from horizontal. | 09-15-2011 |
| 20110220622 | APPARATUS AND METHOD FOR PREFERENTIAL FORMATION OF WELD JOINT - A weld set-up and weld process are provided that alleviate micro-cracking as between a pair of welded pieces. The weld set-up is provided in which the weld pieces are configured to cause micro-cracking to occur in a portion of the welded pieces that is then removed by post-processing machining techniques. This weld set-up includes a primary weld piece and a secondary weld piece with a flange portion that extends over the primary weld piece. When used in conjunction with electron beam welding, application of the weld beam onto the secondary weld piece causes the formation of a preferential nailhead that is formed exclusively in the first weld piece. | 09-15-2011 |
| 20110219743 | INJECTOR ASSEMBLY FOR A ROCKET ENGINE - A cap for a liquid propellant injection assembly of a rocket engine includes a cap body and a valve assembly. The cap body extends between first and second ends. The cap body has a bore that fluidly connects one or more inlets to an outlet. The inlets are disposed in a tubular sidewall of the cap body. The outlet is disposed in the second end of the cap body. The valve assembly includes a valve cap disposed around the first end of the cap body. The valve assembly is adapted to selectively regulate flow of a propellant through the inlets in the cap body as a function of pressure exerted by the propellant against the valve assembly. | 09-15-2011 |
| 20110217464 | METHOD FOR APPLYING A THERMAL BARRIER COATING - A method and apparatus for forming thermally grown alpha alumina oxide scale on a substrate is provided. The method includes the steps of: a) providing a heating chamber having a heat source and an oxidizing gas source selectively operable to provide a stream of oxidizing gas; b) providing at least one substrate disposed in the heating chamber, which substrate has a composition sufficient to permit formation of an alpha alumina scale on one or more surfaces; c) maintaining a vacuum in the heating chamber at a level that inhibits formation of one or more low temperature oxides on the one or more surfaces of the substrate; d) heating at least one of the one or more surfaces of the substrate to a predetermined temperature at or above 1800 degrees Fahrenheit; and e) directing the stream of oxidizing gas at a controlled rate toward one or more heated surfaces of the substrate. | 09-08-2011 |
| 20110214433 | STRAIN TOLERANT BOUND STRUCTURE FOR A GAS TURBINE ENGINE - A gas turbine engine includes bound assemblies with an inner diameter ring, struts, and an outer diameter ring. The strut is connected to the inner diameter ring and extends radially outward therefrom to connect to the outer diameter ring. A strain relief feature is disposed adjacent to or at the connection between the strut and the inner diameter ring and/or the outer diameter ring. The strain relief feature lengthens the arc segment of fillet curvature. For a constant thermal punch load, the lengthened arc segment of fillet curvature results in a decreased maximum strain in the bound assembly. | 09-08-2011 |
| 20110211967 | HYBRID METAL FAN BLADE - A fan blade is disclosed comprising a lightweight metallic airfoil portion and a high-strength sheath portion. The airfoil portion has a forward airfoil edge, a first airfoil surface, and a second airfoil surface. The sheath portion has a sheath head section, a first sheath flank, and a second sheath flank, both flanks extending chordwise from the forward sheath section. The sheath portion is bonded to the airfoil portion such that the sheath head section covers the forward airfoil edge, defining a blade leading edge. The first sheath flank covers a portion of the first airfoil surface proximate the airfoil forward edge, jointly defining a blade suction surface. The second sheath flank covers a portion of the second airfoil surface proximate the airfoil forward edge, jointly defining a blade pressure surface. | 09-01-2011 |
| 20110211965 | HOLLOW FAN BLADE - A blade includes a hollow metal airfoil having an opening to an internal cavity in a first major surface. A metal cover is adhesively bonded to a socket formed around the opening. The cover encloses the cavity and provides a continuous first major surface. | 09-01-2011 |
| 20110206533 | REPAIR OF A COATING ON A TURBINE COMPONENT - A method is provided for repairing an area of damaged coating on a component of a turbine module in a gas turbine engine, the component formed of a base material having a diffusion coating applied to the base material. The repair may be accomplished in place by directly heating the area to which a touch-up coating material has been applied with a hot gas plasma without the need to place the component in an oven for curing and heat treatment of the touch-up coating applied to the damaged area. | 08-25-2011 |
| 20110206519 | FASTENER APERTURE HAVING AN ELONGATED GEOMETRY - An apparatus includes a substrate having one or more fastener apertures that extend therethrough. Each fastener aperture has a centerline and includes first and second circular segmented regions and a central channeled region. Each circular segmented region has a diameter and a segment length that extends along the centerline, wherein the segment length is greater than one-half the diameter. The central channeled region extends along the centerline between the first and the second circular segmented regions. The central channeled region has a height that is less than the diameter of the first and the second circular segmented regions. | 08-25-2011 |
| 20110204162 | ARRANGEMENT FOR CONTROLLING FLUID JETS INJECTED INTO A FLUID STREAM - In an air mixing arrangement wherein a primary fluid is introduced through an opening in a wall to be mixed with a secondary fluid flowing along the wall surface, the opening is airfoil shaped with its leading edge being orientated at an attack angle with respect to the secondary fluid flow stream so as to thereby enhance the penetration and dispersion of the primary fluid stream into the secondary fluid stream. The airfoil shaped opening is selectively positioned such that its angle of attack provides the desired lift to optimize the mixing of the two streams for the particular application. In one embodiment, a collar is provided around the opening to prevent the secondary fluid from contacting the surface of the wall during certain conditions of operation. Multiple openings maybe used such as the combination of a larger airfoil shaped opening with a smaller airfoil shaped opened positioned downstream thereof, or a round shaped opening placed upstream of an airfoil shaped opening. Pairs of openings and associated collars maybe placed in symmetric relationship so as to promote mixing in particular applications, and nozzles maybe placed on the inner side of wall to enhance the flow characteristics of the primary fluid. | 08-25-2011 |
| 20110204122 | HOLLOW STRUCTURES FORMED WITH FRICTION STIR WELDING - A hollow structure, and a method of forming the hollow structure, where the hollow structure includes first and second metal parts, the second metal part having an interior surface and a tapered support member extending from the interior surface. The hollow structure also includes a friction stir welded joint that extends through the first metal part and into the tapered support member. | 08-25-2011 |
| 20110203293 | BEARING COMPARTMENT PRESSURIZATION AND SHAFT VENTILATION SYSTEM - An assembly and method for providing buffer air and/or ventilation air within a gas turbine engine, the assembly includes an accessory gearbox and a centrifugal compressor. The accessory gearbox is connected to a driven shaft of the gas turbine engine. The centrifugal compressor is driven by the accessory gearbox during operation of the gas turbine engine to receive and compress a bleed air to produce the buffer air and/or the ventilation air. The buffer air and/or the ventilation air is communicated within the gas turbine engine. | 08-25-2011 |
| 20110203286 | 3D NON-AXISYMMETRIC COMBUSTOR LINER - A combustor liner with an input end and an output end includes an annular inner wall and an annular outer wall. At least one of the inner wall and outer wall is three-dimensionally contoured. The inner wall and the outer wall form a combustion chamber with the contours creating alternating expanding and constricting regions inside the chamber causing combustion gases to flow in the circumferential and axial directions. | 08-25-2011 |
| 20110200759 | OXIDATION RESISTANT COATINGS, PROCESSES FOR COATING ARTICLES, AND THEIR COATED ARTICLES - A process for applying an oxidation resistant coating to an article includes the steps of mixing at least about 10% by volume to up to about 99% by volume of a slurry at least one silica based material having a viscosity of about 1×10 | 08-18-2011 |
| 20110194941 | CO-CURED SHEATH FOR COMPOSITE BLADE - A method of forming a composite blade having a sheath on a portion of the blade by placing the dry composite blade and sheath in a mold, adding a resin, and curing the resin to integrally bond the blade to the metal sheath. The portion of the blade having the sheath bonded thereto may be at least one of the leading edge, the tip and the trailing edge of the blade. | 08-11-2011 |
| 20110189001 | ROTATABLE VANED NOZZLE FOR A RADIAL INFLOW TURBINE - A variable nozzle system can comprise a gas inlet ring, an opposing gas outlet ring, an actuation ring, guides, and vanes circumferentially spaced about and disposed between the gas inlet ring and the gas outlet ring. The gas inlet ring, the gas outlet ring, and the vanes can form nozzles, the nozzles being variable by rotation of the vanes about a pivot axis. The plurality of guides can extend from the gas inlet ring, the gas outlet ring, or the actuation ring, and the vanes can be connected to the actuation ring, so that each vane can be rotated by rotation of the actuation ring and by sliding against a respective guide from the plurality of guides. The actuation ring can have a gear rack and can be rotated by rotatable engagement of the gear rack with a pinion attached to the end of a rotatable gear shaft. | 08-04-2011 |
| 20110185740 | COMBUSTOR LINER SEGMENT SEAL MEMBER - A combustor for a gas turbine engine is provided that includes a support shell, a forward liner segment, an aft liner segment, and a seal member. The support shell has an interior surface, and exterior surface, and a plurality of impingement apertures disposed within the shell. The forward liner segment and aft liner segment are attached to the inner surface of the shell. The forward liner segment has an edge surface extending between a face surface and a back surface, and a seal shoulder. The aft liner segment has an edge surface extending between a face surface and a back surface, and a seal shoulder. The forward liner segment and the aft liner segment are separated from one another by a gap. The seal member is disposed within the gap. At least some of the plurality of impingement apertures disposed within the shell are aligned with the seal member and oriented to direct cooling air to impinge on the seal member. | 08-04-2011 |
| 20110185737 | COMBUSTOR LINER SEGMENT SEAL MEMBER - A combustor liner segment seal member is provided that includes a center section, a forward flange, and an aft flange. The center section includes a base surface, a gas path surface, a forward side surface, and an aft side surface. The forward flange extends outwardly from the forward side surface, and includes a width, a height, a shell side surface, and a liner side surface. The aft flange extends outwardly from the aft side surface, and includes a width, a height, a shell side surface, and a liner side surface. | 08-04-2011 |
| 20110185736 | GAS TURBINE COMBUSTOR WITH VARIABLE AIRFLOW - An annular combustor and a method for operating a gas turbine engine over a power demand range facilitate combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. The flow primary combustion air admitted into the primary combustion zone is varied in response to power demand from a maximum air flow rate of high power demand to a minimum flow air rate of low power demand, while the flow of dilution air into a quench zone downstream of the primary combustion zone is increased from a minimum air flow rate at high power demand to a maximum air flow rate at low power demand. | 08-04-2011 |
| 20110185735 | GAS TURBINE COMBUSTOR WITH STAGED COMBUSTION - An annular combustor for a gas turbine engine is provided that facilitates staged combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. A method is also provided for operating a gas turbine engine over a power demand range that facilitates staged combustion in a lean direct ignition (LDI) mode over an extended range of operating fuel air ratios. | 08-04-2011 |
| 20110185734 | SYSTEM AND METHOD FOR EQUILIBRATING AN ORGANIC RANKINE CYCLE - Embodiments of an ORC system can be configured to reduce ingress of contaminants from the ambient environment. In one embodiment, the ORC system can comprise a pressure equilibrating unit that comprises a variable volume device for holding a working fluid. The variable volume device can be fluidly coupled to a condenser so that working fluid can move amongst the condenser and the variable volume device. This movement can occur in response to changes in the pressure of the working fluid in the ORC system, and in one example the working fluid is allowed to move when the pressure deviates from atmospheric pressure. | 08-04-2011 |
| 20110185733 | SYSTEM AND METHOD FOR EQUILIBRATING AN ORGANIC RANKINE CYCLE - Embodiments of an ORC system can be configured to reduce ingress of contaminants from the ambient environment. In one embodiment, the ORC system can comprise a pressure equilibrating unit that comprises a variable volume device for holding a working fluid. The variable volume device can be fluidly coupled to a condenser so that working fluid can move amongst the condenser and the variable volume device. This movement can occur in response to changes in the pressure of the working fluid in the ORC system, and in one example the working fluid is allowed to move when the pressure deviates from atmospheric pressure. | 08-04-2011 |
| 20110182743 | THREE-DIMENSIONALLY WOVEN COMPOSITE BLADE WITH SPANWISE WEFT YARNS - A composite blade has a root and a tip in a spanwise direction and a leading edge and a trailing edge in a chordwise direction. The composite blade includes a three-dimensional woven preform having weft yarns and warp yarns. The weft yarns extend in the spanwise direction of the composite blade. The warp yarns interweave the weft yarns and extend in the chordwise direction of the blade. | 07-28-2011 |
| 20110182741 | COMPOSITE FAN BLADE LEADING EDGE RECAMBER - A method of forming a vibration stable airfoil by forming a composite blade having a metallic sheath thereon. The sheath has a head section extending out from the blade by a sufficient distance to permit deformation of the head section. The airfoil is tested to determine the vibrational stability thereof; and the head section is re-cambering to adjust the vibrational stability to a desired level. | 07-28-2011 |
| 20110182740 | FAN AIRFOIL SHEATH - A sheath for a fan airfoil having a leading edge, a trailing edge, a tip, a root, a suction side and a pressure side includes a solid portion to wrap around the airfoil lead edge and over the airfoil tip and means for securing the solid portion to the airfoil. | 07-28-2011 |
| 20110182726 | AS-CAST SHROUD SLOTS WITH PRE-SWIRLED LEAKAGE - A system for forming a seal slot in a shroud assembly has a first element having a longitudinally extending slot and a second element having a longitudinally extending slot for receiving a portion of the first element. The first element has a pair of opposed surfaces and the second element has a pair of opposed surfaces which are oriented at an angle to the surfaces of the first element. The second element may have a planar portion and a curved portion adjacent the planar portion for forming a slot which allows a flow of fluid to exit the slot at a desired angle. | 07-28-2011 |
| 20110180620 | ELIMINATION OF UNFAVORABLE OUTFLOW MARGIN - A panel for use in a gas turbine engine exhaust case is disclosed. The panel has an airfoil section and a flow diverting structure adjacent a leading edge, wherein the flow diverting structure directs fluid flow into an area of the airfoil that lacks sufficient internal pressure for cooling fluid flow. | 07-28-2011 |
| 20110180521 | DEPTH AND BREAKTHROUGH DETECTION FOR LASER MACHINING - A system comprises a working laser beam, a sensing laser beam, first and second optical elements, an optical sensor, an aperture and a controller. The first optical element generates a coaxial beam from the working laser beam and the sensing laser beam. The second optical element focuses the coaxial beam onto a workpiece, such that the working laser beam machines the workpiece and the sensing laser beam reflects from the workpiece. The optical sensor senses an intensity of the reflected sensing beam. The aperture determines a focus position by translating along the reflected sensing beam, such that the reflected intensity is maximized. The controller determining a machining parameter of the working laser beam, based on the focus position. | 07-28-2011 |
| 20110180199 | Powder -metallurgy braze preform and method of use - A method of fabricating and using a braze preform, the method comprising cold pressing a powdered brazing material in a die, thereby forming a braze preform in the die. The braze preform can then be diffusion brazed to a damaged component to repair the component. | 07-28-2011 |
| 20110179763 | Particle separator and debris control system - A gas turbine engine system includes an air inlet, an inlet particle separator located at the air inlet and having a blower selectively driven by a variable output motor, and a controller for dynamically controlling the variable output motor that selectively drives the blower of the inlet particle separator. | 07-28-2011 |
| 20110176927 | COMPOSITE FAN BLADE - A composite fan blade includes a first filament reinforced airfoil ply section, a second filament reinforced airfoil ply section and a three dimensionally woven core. The woven core is located between the first and second reinforced airfoil ply sections. The woven core includes first yarns extending in a chordwise direction and second yarns extending in a spanwise direction. There are a greater number of second yarns in a first region of the woven core than in a second region of the woven core. | 07-21-2011 |
| 20110167635 | METHOD OF REPAIRING SHROUDED TURBINE BLADES WITH CRACKS IN THE VICINITY OF THE OUTER SHROUD NOTCH - A method of repairing a damaged notch fillet radius of a turbine blade shroud includes a replacement notch fillet radius and a hardface nugget. The damaged notch fillet radius is blended out and a hardface structure positioned proximate the damaged notch fillet radius is removed. A replacement notch fillet radius is formed with weld filler and the hardface nugget is welded to at least a portion of the replacement notch fillet radius. | 07-14-2011 |
| 20110167634 | COATED GAS TURBINE ENGINE COMPONENT REPAIR - A method of repairing a component of a gas turbine engine that includes a metallic substrate, an existing aluminide coating, and a diffusion layer formed in the metallic substrate adjacent to the coating. The method includes removing at least a portion of the existing aluminide coating, removing material forming the diffusion layer, applying a new metallic layer to the metallic substrate, and applying a new aluminide coating over the new metallic layer to form a new diffusion layer in the new metallic layer. The new metallic layer is a substantially homogeneous material that is substantially similar in chemical composition to that of the metallic substrate, and the new metallic layer forms a structural layer having a thickness selected to provide a specified contour to the component. | 07-14-2011 |
| 20110165335 | DISPERSION STRENGTHENED RARE EARTH STABILIZED ZIRCONIA - A process for forming a coating on a substrate, such as a turbine engine component, is provided. The process comprises the steps of: providing a first rare earth oxide stabilized zirconia composition; providing a second composition selected from the group consisting of a yttria stabilized zirconia composition, a ceria stabilized zirconia composition, alumina, a chromia-alumina composition, a gadolinia stabilized zirconia composition, and mixtures thereof; blending the first rare earth oxide stabilized zirconia composition with the second composition to form a blended powder; and depositing the blended powder onto the substrate. Articles having the coating formed from the above process are also described. | 07-07-2011 |
| 20110155508 | ON-DEMAND LUBRICATION SYSTEM FOR IMPROVED FLOW MANAGEMENT AND CONTAINMENT - An oil circulation system for a gas turbine engine includes a variable displacement scavenge pump for pumping oil, a first actuator for controlling displacement of the scavenge pump, a variable displacement pressure pump for pumping oil, a second actuator for controlling displacement of the pressure pump, and a drive shaft. Both the scavenge pump and the pressure pump are driven by the drive shaft. | 06-30-2011 |
| 20110153295 | METHOD AND SYSTEM FOR MODELING THE PERFORMANCE OF A GAS TURBINE ENGINE - A method for modeling the performance of a gas turbine engine is provided. The method includes the steps of: 1) providing a processor; 2) inputting flight condition parameter data and engine output parameter data into a gas turbine engine model operating on the processor, which model includes a physics-based engine model that uses the flight condition parameter data to produce estimated engine output parameter data, and determines residuals from the engine output parameter data and the estimated engine output parameter data; 3) partitioning the flight condition parameter data and residuals into training data and testing data; 4) performing a correlation reduction on the training data, which analysis produces correlation adjusted training data; 5) performing an orientation reduction on the correlation adjusted training data, which reduction produces orientation adjusted training data; 6) reviewing the orientation adjusted training data relative to at least one predetermined criteria, and iteratively repeating the steps of performing a correlation reduction and an orientation reduction using the orientation adjusted training data if the criteria is not satisfied, and if the criteria is satisfied outputting the orientation adjusted training data; 7) producing estimated corrections to the orientation adjusted training data using one or more neural networks; 8) evaluating the neural adjusted data using the partitioned testing data; and 9) modeling the performance of the gas turbine using the estimated corrections to the orientation adjusted training data. | 06-23-2011 |
| 20110150636 | IN-SITU TURBINE BLADE TIP REPAIR - The tip portion of a blade of a gas turbine is repaired in-situ in the gas turbine without removing the blade from the turbine rotor disk. To facilitate the in-situ repair of the tip portion of the turbine blades, one or more access holes may be provided in the turbine shroud or blade outer air seal circumscribing the plurality of turbine blades extending radially outward from the turbine rotor dick. | 06-23-2011 |
| 20110146298 | RETAINING MEMBER FOR USE WITH GAS TURBINE ENGINE SHAFT AND METHOD OF ASSEMBLY - A gas turbine engine assembly includes a rotor shaft having a threaded portion, a rotor stack support positioned radially outward from the rotor shaft, and a locking nut. The locking nut includes a body portion, a threaded portion located on a radially inner surface of the body portion, and a lug portion extending from the body portion. The rotor stack support radially rests upon the body portion of the locking nut, and the rotor stack support axially rests against the lug portion of the locking nut. The threaded portion of the locking nut and the threaded portion of the rotor shaft are engaged together. | 06-23-2011 |
| 20110143164 | LOW SULFUR NICKEL BASE SUBSTRATE ALLOY AND OVERLAY COATING SYSTEM - A coated article having an improved coating oxidation life includes a superalloy substrate material having a composition which includes sulfur, herein the sulfur is present in an amount less than 1 ppm; and an overlay coating formed over a surface of the substrate material. | 06-16-2011 |
| 20110143043 | PLASMA APPLICATION OF THERMAL BARRIER COATINGS WITH REDUCED THERMAL CONDUCTIVITY ON COMBUSTOR HARDWARE - A process for forming a thermal barrier coating comprises the steps of providing a substrate, providing a gadolinia stabilized zirconia powder, and forming a thermal barrier coating having at least one of a porosity in a range of from 5 to 20% and a dense segmented structure on said substrate by supplying the gadolinia stabilized powder to a spray gun and using an air plasma spray technique. | 06-16-2011 |
| 20110139393 | PROCESS TO CAST SEAL SLOTS IN TURBINE VANE SHROUDS - A process for casting a turbine engine component is provided. The process comprises the steps of placing a refractory core assembly comprising two intersecting plates in a die, encapsulating the refractory core assembly in a wax pattern having the form of the turbine engine component, forming a ceramic shell mold about the wax pattern, removing the wax pattern, and pouring molten material into the ceramic shell mold to form the turbine engine component. | 06-16-2011 |
| 20110138809 | OPERATING A SUB-SEA ORGANIC RANKINE CYCLE (ORC) SYSTEM USING INDIVIDUAL PRESSURE VESSELS - A method and system for generating electrical power for sub-sea applications includes assembling each of the main components ( | 06-16-2011 |
| 20110138769 | FAN CONTAINMENT CASE - A fan section of a gas turbine engine having an axially extending centerline is provided. The fan section includes a fan assembly, a fan containment case, and a plurality of structural exit guide vanes. The fan assembly includes a fan rotor hub centered on and rotatable about the centerline, and a plurality of fan blades that are attached to and extend radially out from the fan rotor hub. The fan containment case is disposed radially outside of and circumferentially around the fan assembly. The fan containment case includes a shell and an integral flange ring that extends around the circumference of the shell, a circumferentially extending blade outer air seal disposed between the fan blades and the shell. The structural exit guide vanes are mechanically attached to the fan containment case at a position aligned with the integral flange ring. | 06-16-2011 |
| 20110138625 | REPAIR OF INTEGRALLY BLADED ROTORS - A method is provided for repairing a damaged rotor blade on an integrally bladed rotor by removing a damaged portion of a damaged blade leaving a blade stub extending outwardly from the disk and performing a linear friction welding operation to attach a replacement blade segment to the blade stub. The rotor may be disposed operation using a linear friction welding apparatus. The method includes disposing a support collar about the blade stub and securing the support collar to the linear friction welding apparatus prior to a commencement of the bonding operation. A lower surface of the support collar is contoured to mate with a portion of an outer circumference surface of the rotor disk. | 06-16-2011 |
| 20110138574 | HINGE ARM REPAIR - A hinge arm having a hinge pin hole having a worn bore, for example a hinge arm of a hinge assembly of an access door on a gas turbine engine, is repaired by forming in the worn hinge arm a new hole with a keyway extending outwardly from the new hole. A keyed bushing is provided having a cylindrical body having a central bore extending axially therethrough and a key extending outwardly from the cylindrical hole. The keyed bushing is inserted into the new hole and keyway formed in the hinge arm thereby preventing rotation of the bushing relative to the hinge arm. | 06-16-2011 |
| 20110135482 | TIP VORTEX CONTROL - A rotor blade for a gas turbine engine includes an attachment and an airfoil. The airfoil has a stagger angle, a base region, a transition region and a tip region. The stagger angle changes as the airfoil extends between the attachment and a tip. The base region is disposed adjacent to the attachment. The transition region is located between the base and the tip regions. A rate of the change of the stagger angle in the transition region is greater than a rate of the change of the stagger angle in the base region. The rate of the change of the stagger angle in the transition region is greater than a rate of change of the stagger angle in the tip region. | 06-09-2011 |
| 20110135446 | Castings, Casting Cores, and Methods - If a refractory metal core (RMC) is punched, the punching asymmetry may be reflected in an asymmetry of the cast article features cast by the punched features. The punched features may have a shear zone and a fracture zone. The shear zone of the RMC will cast a relatively narrow portion of the post near one end; whereas the fracture zone will cast a relatively broader portion near the other end. The broader portion will also have a relatively shallow transition to the adjacent face of the slot-like passageway cast by the RMC. Where there is a stress asymmetry in the cast article in-use, the punching direction may be chosen so that the relatively broad portions of the post fall along the relatively higher stress face of the passageway. | 06-09-2011 |
| 20110132876 | ARTICLE HAVING DIFFUSER HOLES AND METHOD OF MAKING SAME - The diffusion opening in the cooling hole of an airfoil is formed by an EDM process in which the outwardly flaring sidewalls of the opening, rather then having surfaces that are approximated to be smooth by having many small ribs formed therein, are formed with a relatively few ribs with both longitudinally extending and radially extending surfaces that are substantially greater in dimension than those as normally formed. In this manner, the machining process is simplified and expedited, while at the same time, the cooling efficiency is increased. | 06-09-2011 |
| 20110130246 | MOUNTING SYSTEM FOR A PLANATARY GEAR TRAIN IN A GAS TURBINE ENGINE - A mounting system for a planetary gear train in a gas turbine engine comprises a support strut, a deflection flange and a deflection limiter. The support strut extends between a stationary engine case and a rotating engine shaft that provides input to the planetary gear train in the gas turbine engine. The deflection flange extends from a rotating output component of the planetary gear train. The deflection limiter is connected to the support strut and engages the deflection flange when the gear train becomes radially displaced. | 06-02-2011 |
| 20110129348 | CORE DRIVEN PLY SHAPE COMPOSITE FAN BLADE AND METHOD OF MAKING - A method of forming a composite airfoil having a suction side and pressure side includes the steps of designing a mold, designing a woven core, designing a plurality of plies and assembling the designed mold, core and plies to create the composite airfoil. The hollow mold has an inner surface which defines the surface profile of the composite airfoil. The plurality of plies is designed to fit between the inner surface of the mold and an outer surface of the woven core. The plurality of plies is designed after the step of designing the woven core. | 06-02-2011 |
| 20110126543 | COMBUSTOR PANEL ARRANGEMENT - A combustor module for a gas turbine engine is provided that includes a first annular liner assembly extending along a longitudinal axis of the engine. The first annular liner assembly includes a first annular support shell and a plurality of first heat shield panels coupled to the first annular support shell. The first heat shield panels form a segmented ring defining a plurality of first axial seams therebetween. The combustor module further includes a bulkhead coupled to the first annular liner assembly. The bulkhead provides a plurality of fuel nozzles for passing a first mass flow comprising fuel and air. The combustor module further includes a second annular liner assembly coupled to the bulkhead. The second annular liner assembly is in spaced-apart generally coaxial relationship from the first annular liner assembly by a channel height H. The second annular liner assembly includes an air admittance hole having a mean diameter D extending along a hole axis. The hole axis is offset from the first axial seam defined by the first heat shield panels. | 06-02-2011 |
| 20110126539 | COMBINED GEOTHERMAL AND SOLAR THERMAL ORGANIC RANKINE CYCLE SYSTEM - In a system where the thermal energy of a geothermal fluid is applied to an ORC system, the energy is enhanced by the use of solar energy to thereby increase the temperature of the fluid being applied by the ORC system. A single heat exchanger version provides for direct heat exchange relationship with the geothermal and solar fluids, whereas a two heat exchanger version provides for each of the geothermal and solar fluids to be in heat exchange relationship with the working medium of the ORC system. Control features are provided to selectively balance the various fluid flows in the system. | 06-02-2011 |
| 20110123326 | BELLOWS PRELOAD AND CENTERING SPRING FOR A FAN DRIVE GEAR SYSTEM - An assembly includes a first tapered roller bearing and a second tapered roller bearing and a bellows spring. The bellows spring is disposed adjacent the first tapered roller bearing and the second tapered roller bearing. | 05-26-2011 |
| 20110123322 | FLOW PASSAGE FOR GAS TURBINE ENGINE - Inefficiencies associated with the formation of horseshoe vortices in a gas turbine engine having a row of radially extending airfoils, are mitigated with an improved flow passage defined by the suction and pressure surface of two adjacent airfoils and endwalls (disposed proximal to the radially inner and outer ends of the airfoils). The flow passage includes a ridge disposed in one of the endwalls adjacent a suction surface of one of the airfoils, the ridge extending longitudinally from a location at or near the maximum circumferential extent of the airfoil suction side to a downstream location proximal to the trailing edge of the suction surface. | 05-26-2011 |
| 20110120154 | PASSIVE OIL LEVEL LIMITER - A vapor expansion/compression system having an oil sump for lubrication of the turbine/compressor bearings includes an oil and/or vapor vent line leading from a strategic location in the oil sump to the condenser such that the oil level in the sump is limited to the predetermined level, and any excess oil is pumped from the sump to thereby prevent the oil level from reaching a level at which bearings failure could be caused due to a high oil level. | 05-26-2011 |
| 20110120134 | GAS TURBINE COMBUSTOR - An annular combustor for a gas turbine has a combustion chamber having an interior volume that, in longitudinal section, includes a forward volume, an intermediate volume and an aft volume. The forward volume represents from about 30% to about 40% of the combustor interior volume, the intermediate volume represents from about 10% to about 20% of the combustor interior volume, and the aft volume represents from about 40% to about 60% of the combustor interior volume. | 05-26-2011 |
| 20110117385 | METHOD OF REPAIR ON NICKEL BASED HPT SHROUDS - A method of repairing an abradable, nickel single crystal material having an abraded surface area by applying a diffusion layer to the abraded surface area comprising a nickel-based alloy having a quantity of an additional element selected from silicon and boron; and also applying an abradable layer to the material comprising a cobalt-based alloy to restore the abraded surface area. | 05-19-2011 |
| 20110110783 | Trunnion hole repair utilizing interference fit inserts - A vane shroud for a gas turbine engine has a cylindrical shaped body and a hole arranged at a circumference of the body. The hole is for receiving a trunnion from a variable vane. A counterbore is arranged concentrically around the hole on a radially inward facing surface of the body, and receives a trunnion hub from the variable vane. A shrink-fit insert is seated into the counterbore such that the trunnion hub is provided with a contact surface, and the shrink-fit insert is secured to the shroud through an interference fit. | 05-12-2011 |
| 20110108360 | OIL CAPTURE AND BYPASS SYSTEM - An assembly for capturing lubricating liquid from a fan drive gear system of a gas turbine engine includes a gutter and an auxiliary reservoir. The gutter is positioned radially outward from a centerline axis of the gear system for capturing lubricating liquid slung from the gear system. The auxiliary reservoir is fluidically connected to the gutter for receiving and storing the lubricating liquid from the gutter. A portion of the auxiliary reservoir and a portion of the gutter have a same radial distance from the centerline axis of the gear system. | 05-12-2011 |
| 20110091346 | Forging deformation of L12 aluminum alloys - A method for producing a high strength aluminum alloy parts containing L1 | 04-21-2011 |
| 20110091345 | Method for fabrication of tubes using rolling and extrusion - A method for producing a high strength aluminum alloy tubing containing L1 | 04-21-2011 |
| 20110088865 | PROCESS TO CAST SEAL SLOTS IN TURBINE VANE SHROUDS - A process for casting a turbine engine component is provided. The process comprises the steps of placing a refractory core assembly comprising two intersecting plates in a die, encapsulating the refractory core assembly in a wax pattern having the form of the turbine engine component, forming a ceramic shell mold about the wax pattern, removing the wax pattern, and pouring molten material into the ceramic shell mold to form the turbine engine component. | 04-21-2011 |
| 20110088510 | Hot and cold rolling high strength L12 aluminum alloys - A method and apparatus for producing high strength aluminum alloys from a powder containing Ll | 04-21-2011 |
| 20110088403 | NOZZLE-AREA RATIO FLOAT BIAS - A mechanism included in a convergent-divergent nozzle connected to an aft portion of a gas turbine engine, which mechanism includes a convergent flap configured to be kinematically connected to the aft portion of the gas turbine engine, a divergent flap pivotally connected to the convergent flap and configured to be kinematically connected to the aft portion of the gas turbine engine, at least one actuator configured to be connected to the aft portion of the gas turbine engine and to mechanically prescribe the position of the convergent flap and the divergent flap by moving through one or more positions, and at least one biasing apparatus. The biasing apparatus is configured to position at least one of the convergent flap and the divergent flap independent of the positions of the actuator by substantially balancing a force on the at least one of the convergent flap and the divergent flap produced by a gas pressure on the nozzle. | 04-21-2011 |
| 20110086179 | THERMAL BARRIER COATING WITH A PLASMA SPRAY TOP LAYER - A turbine engine component has a substrate, a thermal barrier coating deposited onto the substrate, and a sealing layer of ceramic material on an outer surface of the thermal barrier coating for limiting molten sand penetration. | 04-14-2011 |
| 20110085932 | METHOD OF FORMING HIGH STRENGTH ALUMINUM ALLOY PARTS CONTAINING L12 INTERMETALLIC DISPERSOIDS BY RING ROLLING - A method and apparatus produces high strength aluminum alloy parts from a powder containing L1 | 04-14-2011 |
| 20110077783 | SYSTEM AND METHOD FOR DESIGN AND CONTROL OF ENGINEERING SYSTEMS UTILIZING COMPONENT-LEVEL DYNAMIC MATHEMATICAL MODEL - A control system comprises a controller for positioning an actuator in a working fluid flow and a model processor for directing the controller as a function of a model feedback. The model processor comprises an output module, a comparator and an estimator. The output module generates the model feedback as a function of a constraint, a model state and a model input describing fluid parameters measured along the working fluid flow. The comparator generates an error by comparing the model feedback to the model input. The estimator generates the constraint and the model state, such that the error is minimized. | 03-31-2011 |
| 20110076405 | HOLE DRILLING WITH CLOSE PROXIMITY BACKWALL - A method of making a coated component with a close proximity backwall is achieved by applying a coating to a pre-existing workpiece that contains a substrate with a plurality of apertures. The substrate is in close proximity to a backwall. The coating is removed from the plurality of apertures with a fluid jet cutting system. The fluid jet cutting system has a fluid jet that does not include a particulate material additive. | 03-31-2011 |
| 20110073251 | COMPONENT BONDING PREPARATION METHOD - A component bonding preparation method employs plasma to prepare a component for bonding. A first plasma is used to hydroxylate a component surface. A second plasma is used to silanize the component surface. | 03-31-2011 |
| 20110071014 | HYBRED POLYMER CVI COMPOSITES - A method of forming a highly densified chemical matrix composite CMC from a preform of a matrix of a non-oxide ceramic and continuous ceramic fibers. An interface coating is added, followed by partially densifying the preform with a resin to increase the density of the preform using a polymer infiltration pyrolysis PIP) process one or more times. A chemical vapor infiltration (CVI) process is used to bring the CMC to a final desired density. | 03-24-2011 |
| 20110071013 | CERAMIC MATRIX COMPOSITE SYSTEM AND METHOD OF MANUFACTURE - A ceramic matrix composite is formed from a non-oxide ceramic and continuous ceramic fibers. The matrix includes a hafnium donor in the matrix in an amount sufficient to harden the composite at elevated temperatures to prevent ablation of the composite. The matrix also includes a boron donor in the matrix in an amount sufficient to lower the glass transition temperature of the composite to flow over cracks formed in the composite. The method to form the matrix is selected from polymer infiltration pyrolysis, chemical vapor infiltration, and sequential polymer infiltration pyrolysis and chemical vapor infiltration. | 03-24-2011 |
| 20110068010 | ANODE MEDIA FOR USE IN ELECTROPLATING PROCESSES, AND METHODS OF CLEANING THEREOF - A method for cleaning anode media, the method comprising removing the anode media from an electroplating system, and removing scale coatings from substrates of the anode media by vibrational polishing the anode media with abrasive particles. | 03-24-2011 |
| 20110064599 | DIRECT EXTRUSION OF SHAPES WITH L12 ALUMINUM ALLOYS - A method for producing a high strength aluminum alloy brackets, cases, tubes, ducts, beams, spars and other parts containing L1 | 03-17-2011 |
| 20110064580 | TURBOFAN FLOW PATH TRENCHES - An integrally bladed disk includes a rim and blades. The blades are circumferentially spaced and extend integrally outward from the rim to define a flow path therebetween. The periphery of the rim forms a flow surface and includes a trench that is disposed adjacent the blades. | 03-17-2011 |
| 20110064515 | ANCHOR AND SPLICE PLATE ASSEMBLY FOR AXIALLY SPLIT COMPOSITE DUCT OR PRESSURE VESSEL - An assembly includes first and second shell members. An anchor plate includes a body portion, first and second arrays of fastener openings, and anchor plate attachment features extending from a first side of the body portion and each having a fastener opening. Inserts are positioned within each fastener opening of the first array and the second array. A first set of fasteners is secured through the first array of fastener openings in the anchor plate, the first shell member and the splice plate and engaged with a corresponding one of the inserts. A second set of fasteners is secured through the second array of fastener openings in the anchor plate, the second shell member and the splice plate and engaged with a corresponding one of the inserts. A third set of fasteners is secured through the fastener openings of the anchor plate attachment features and the first shell member. | 03-17-2011 |
| 20110061767 | COMPONENT REMOVAL TOOL AND METHOD - An apparatus includes a carriage, a first tooth, a second tooth, a feed block and a drive element. The first and second teeth are shaped for breaking bonds and attached to a first side of the carriage. The first and second teeth are spaced apart and generally parallel to each other. The drive element moves the carriage away from the feed block so that the first and second teeth break bonds when they encounter them. A method for removing a bonded component from a structure includes positioning a tool having first and second teeth between the component and the structure. The teeth are positioned between trailing edges of the component and the structure. The method includes driving the teeth from the trailing edges of the structure towards a leading edge of the structure to break bonds between the component and the structure so that the component can be removed. | 03-17-2011 |
| 20110061494 | SUPERPLASTIC FORMING HIGH STRENGTH L12 ALUMINUM ALLOYS - A method and apparatus produces high strength aluminum alloys from a powder containing L1 | 03-17-2011 |
| 20110061362 | ARRANGEMENT FOR CONTROLLING FLUID JETS INJECTED INTO A FLUID STREAM - In an air mixing arrangement wherein a primary fluid is introduced through an opening in a wall to be mixed with a secondary fluid flowing along the wall surface, the opening is airfoil shaped with its leading edge being orientated at an attack angle with respect to the secondary fluid flow stream so as to thereby enhance the penetration and dispersion of the primary fluid stream into the secondary fluid stream. The airfoil shaped opening is selectively positioned such that its angle of attack provides the desired lift to optimize the mixing of the two streams for the particular application. In one embodiment, a collar is provided around the opening to prevent the secondary fluid from contacting the surface of the wall during certain conditions of operation. Multiple openings maybe used such as the combination of a larger airfoil shaped opening with a smaller airfoil shaped opened positioned downstream thereof, or a round shaped opening placed upstream of an airfoil shaped opening. Pairs of openings and associated collars maybe placed in symmetric relationship so as to promote mixing in particular applications, and nozzles maybe placed on the inner side of wall to enhance the flow characteristics of the primary fluid. | 03-17-2011 |
| 20110056221 | ACTIVE STRESS CONTROL DURING RAPID SHUT DOWN - A closed loop refrigerant expansion system with a tube and shell condenser is provided with a control which, upon shutdown, causes the flow of refrigerant to reverse from the evaporator to the condenser to thereby both reduce the amount of refrigerant vapor passing to the condenser and increase the amount of liquid refrigerant in the condenser to thereby reduce the maximum temperature load in the condenser. Reverse flow can be made to occur either by reversing the direction of the refrigerant pump or opening a bypass valve around the pump. | 03-10-2011 |
| 20110056213 | SPOOL SUPPORT STRUCTURE FOR A MULTI-SPOOL GAS TURBINE ENGINE - A gas turbine engine is provided that includes a low pressure spool, a high pressure spool, a stationary support frame, and at least one support arch. The low pressure spool extends between a low pressure compressor and a low pressure turbine. The high pressure spool extends between a high pressure compressor and a high pressure turbine. The spools are rotatable about a center axis of the engine. The support arch has a stationary support mount disposed between a low spool mount and a high spool mount. The support arch is disposed relative to the spools and the stationary support frame so that a load from each spool caused by the rotation of that spool can be transferred to the stationary support frame through the support arch. The support arch can freely rotate about the center axis of the engine relative to the spools and the stationary structural frame. | 03-10-2011 |
| 20110056208 | REVERSED-FLOW CORE FOR A TURBOFAN WITH A FAN DRIVE GEAR SYSTEM - A gas turbine engine comprises a fan drive gear system, a low spool connected to the fan drive gear system, and a high spool disposed aft of the low spool. The low spool comprises a rearward-flow low pressure compressor disposed aft of the fan drive gear system, and a forward-flow low pressure turbine disposed aft of the low pressure compressor. The high spool comprises a forward-flow high pressure turbine disposed aft of the low pressure turbine, a combustor disposed aft of the high pressure turbine, and a forward-flow high pressure compressor disposed aft of the combustor. | 03-10-2011 |
| 20110054704 | HIGH FIDELITY INTEGRATED HEAT TRANSFER AND CLEARANCE IN COMPONENT-LEVEL DYNAMIC TURBINE SYSTEM CONTROL - A system comprises a rotary apparatus, a control law and a processor. The rotary apparatus comprises a rotor and a housing forming a gas path therebetween, and the control law controls flow along the gas path. The processor comprises an output module, a plurality of temperature modules, a thermodynamic module, a comparator and an estimator. The output module generates an output signal as a function of a plurality of rotor and housing temperatures defined along the gas path, and the temperature modules determine time derivatives of the rotor and housing temperatures. The thermodynamic module models boundary conditions for the gas path, and the comparator determines errors in the boundary conditions. The estimator estimates the rotor and housing temperatures based on the time derivatives, such that the errors are minimized and the flow is controlled. | 03-03-2011 |
| 20110053468 | THERMAL MECHANICAL SKIVE FOR COMPOSITE MACHINING - An apparatus for thermal mechanical machining of composite materials includes a head, a drive, and a shaft. The head has an abrasive face. The drive is coupled to the apparatus to move the head to produce abrasion of the composite material by the abrasive face. The shaft includes a passageway that communicates a heated gas to an interface between the abrasive face and the composite material. The gas removes particles that result from the abrasion of the composite material by the abrasive face. The gas can be heated sufficiently to carbonize or vaporize organic constituents of the composite material. | 03-03-2011 |
| 20110052932 | FABRICATION OF L12 ALUMINUM ALLOY TANKS AND OTHER VESSELS BY ROLL FORMING, SPIN FORMING, AND FRICTION STIR WELDING - A method for producing high strength aluminum alloy tanks and other vessels containing L1 | 03-03-2011 |
| 20110052405 | COMPOSITE AIRFOIL WITH LOCALLY REINFORCED TIP REGION - A composite airfoil has a root, a tip, a root region and a tip region. The composite airfoil further includes a woven core, a first filament reinforced airfoil ply, a second filament reinforced airfoil ply and a local reinforcement laminate section. The woven core extends from the root to the tip of the composite airfoil. The first filament reinforced airfoil ply is stacked on the woven core and the second filament reinforced airfoil ply is stacked adjacent to the first filament reinforced airfoil ply on the woven core. The local reinforcement laminate section is at the tip region of the composite airfoil and comprises a first reinforcement ply that does not extend to the root region. The local reinforcement laminate section increases a chordwise flexural stiffness of the tip region. | 03-03-2011 |
| 20110052384 | CERAMIC TURBINE SHROUD SUPPORT - A turbine shroud includes a ceramic shroud ring configured to surround a plurality of turbine rotor blades, a plurality of slots circumferentially distributed around the ceramic shroud ring, a forward metallic support ring, a plurality of tabs attached to a forward edge of the forward metallic support ring configured to engage a turbine support case, and a plurality of tabs attached to an aft edge of the forward metallic support ring received by the slots of the ceramic shroud ring. Only two axially extending radial surfaces of each of the tabs attached to the aft edge of the forward metallic support ring are configured to contact the slots of the ceramic shroud ring. | 03-03-2011 |
| 20110052370 | ROBUST FLOW PARAMETER MODEL FOR COMPONENT-LEVEL DYNAMIC TURBINE SYSTEM CONTROL - A system comprises an apparatus, an actuator and a processor. The apparatus defines a flow path through an aperture, the aperture defines a pressure drop along the flow path, and the actuator regulates fluid flow across the pressure drop. The processor comprises a flow module, a comparator, an estimator and a control law. The flow module maps a flow curve relating a flow parameter to a pressure ratio, and defines a solution point located on the flow curve and a focus point located off the flow curve. The comparator generates an error as a function of a slope defined between the focus and solution points. The estimator moves the solution point along the flow curve, such that the error is minimized. The control law directs the actuator to position the control element, such that the flow parameter describes the fluid flow and the pressure ratio describes the pressure drop. | 03-03-2011 |
| 20110048111 | METHOD OF LEAK TESTING AEROSPACE COMPONENTS - A method includes providing an aerospace electrical component having a wall and a feature that extends through the wall, introducing a detectable residue-free gas on a first side of the wall, and testing for the presence of the detectable gas on a second side of the wall. | 03-03-2011 |
| 20110048024 | GAS TURBINE COMBUSTOR WITH QUENCH WAKE CONTROL - A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber and includes a circumferential row of a plurality of relatively large combustion dilution air admission holes and a circumferential row of a plurality of smaller quench air admission holes disposed downstream with respect to the flow of combustion gas products. The plurality of quench air admission holes are arranged with respect to the plurality of relatively large dilution air admission holes disposed upstream thereof such that there is associated with each dilution air admission hole a first quench air admission hole and a second quench air hole, the first quench air hole being offset laterally in a first lateral direction and the second quench air hole being offset laterally in a second lateral direction opposite to the first direction. | 03-03-2011 |
| 20110047959 | AIR PARTICLE SEPARATOR FOR A GAS TURBINE ENGINE - An assembly for separating debris from a pressurization air flow adjacent to a bearing compartment seal includes a baffle and a capture annulus. The baffle is disposed in the pressurization air flow upstream of the compartment seal. The capture annulus is positioned adjacent the baffle and has a cavity to collect the debris. | 03-03-2011 |
| 20110045181 | APPLYING VAPOUR PHASE ALUMINIDE COATING ON AIRFOIL INTERNAL CAVITIES USING IMPROVED METHOD - A method of applying an aluminide coating to an internal cavity of a metal part, such as an airfoil. The method includes inserting a perforated sealed container containing a source of aluminide into the internal cavity and vaporizing the aluminide source to form an aluminide coating on the wall surface of the cavity. | 02-24-2011 |
| 20110044844 | HOT COMPACTION AND EXTRUSION OF L12 ALUMINUM ALLOYS - A method and apparatus produces high strength aluminum alloys from a powder containing L1 | 02-24-2011 |
| 20110041963 | HEAT TREATABLE L12 ALUMINUM ALLOYS - A method of forming high temperature heat treatable aluminum alloys that can be used at temperatures from about −420° F. (−251° C.) up to about 650° F. (343° C.) are described. The alloys are strengthened by dispersion of particles based on the L1 | 02-24-2011 |
| 20110033284 | STRUCTURALLY DIVERSE THERMAL BARRIER COATINGS - A coated article includes an article having at least one surface and a thermal barrier coating system disposed upon the at least one surface. The thermal barrier coating system has at least two layers, with each layer having a different microstructure. The microstructure of each layer may be any one of the following: columnar, amorphous, randomized, and splat-like. The thermal barrier coating system typically exhibits a thermal conductivity of no more than 16 BTU in/hr ft | 02-10-2011 |
| 20110031223 | INSERT WELD REPAIR - A method of repairing an aperture and adjacent defect wherein one or more defects proximate an aperture are removed from the base material proximate the aperture in an arc shape. An insert containing a profile that corresponds to the arc shaped section of the base material removed is welded to the base material. The weld is accomplished with a single curved path. Next, the insert is removed to create a repaired aperture and a surrounding section that has essentially the same dimensions as the aperture and surrounding section prior to the repair. | 02-10-2011 |
| 20110030387 | MID-TURBINE FRAME TORQUE BOX HAVING A CONCAVE SURFACE - A single point load structure transfers a first load from a first bearing cone and a second load from a second bearing cone of a gas turbine engine to a plurality of struts. The single point load structure includes a stem, a branch connected to the stem and a torque box connected to the plurality of struts for absorbing the first and second loads from the stem and the branch. The stem has a concave surface that opens in a radially outward direction with respect to a rotational axis of the gas turbine engine. | 02-10-2011 |
| 20110030386 | MID-TURBINE FRAME - A mid-turbine frame connected to at least one mount of a gas turbine engine transfers a first load from a first bearing and a second load from a second bearing to the mount. The mid-turbine frame includes a single point load shell structure and a plurality of struts. The single point load shell structure combines the first load and the second load into a combined load. The plurality of struts is connected to the single point load structure and transfers the combined load from the single point load shell structure to the mount. | 02-10-2011 |
| 20110027573 | Lubricated Abradable Coating - A coated article has a metallic substrate. A coating layer comprises at least 33% by volume ceramic, at least 5% by volume of a lubricant selected from the group consisting of hexagonal boron nitride (hBN), clay, graphite, and mixtures thereof. The coating layer lacks a metal phase of more than 5% by volume. | 02-03-2011 |
| 20110027569 | Fastener - A fastener has a metallic member and a nonmetallic cap member. The metallic member has a head and a shaft. The shaft has a first end at the head, a distal second end, and an externally threaded portion. The metallic member further has a projection extending from the head opposite the shaft. The cap member is mated to the head opposite the shaft and retained by the projection. | 02-03-2011 |
| 20110024398 | INSERT WELD REPAIR - A method of repairing an aperture and adjacent defect in a part which is started by removing one or more defects adjacent an aperture in a base material. The material is removed to create a weld seam that extends past an area of high stress concentration on the aperture. An insert of material containing a profile that corresponds to the profile of the base material removed adjacent the aperture and a combination top and runoff plate that encompasses the insert of material are provided. A backing plate is inserted underneath the combination top and runoff plate and insert such that there remains an air space between the backing plate and the combination plate which prevents the combination plate from becoming fused to the backing plate during a welding process. The insert is welded to the base material, and the backing plate is removed. Excess material is removed from the insert to obtain an aperture containing a profile essentially the same as the profile of the aperture prior to initiating the repair. | 02-03-2011 |
| 20110020115 | REFRACTORY METAL CORE INTEGRALLY CAST EXIT TRENCH - A refractory metal core for use in casting a turbine engine part has a main portion and a plurality of tabs extending from the main portion. An end portion having a first edge which joins together the plurality of tabs and a second edge, opposed to the first edge, located internally of an exterior boundary of the cast part may be provided. | 01-27-2011 |
| 20110017359 | HIGH STRENGTH L12 ALUMINUM ALLOYS - High strength heat treatable aluminum alloys that can be used at temperatures from about −420° F. (−251° C.) up to about 650° F. (343° C.) are described. The alloys are strengthened by dispersion of particles based on the L1 | 01-27-2011 |
| 20110008125 | THREADED FLANGED BUSHING FOR FASTENING APPLICATIONS - A fastening system includes a first material having a first hole, a second material having a second hole, a bushing, and a threaded bolt. The bushing is located in the second hole and has a first end abutting the first material and a flanged second for securing the bushing to the second material. A threaded bolt extends through the first hole and into the bushing to secure the first material flush against the second material. | 01-13-2011 |
| 20110005231 | FUEL NOZZLE GUIDE PLATE MISTAKE PROOFING - A fuel nozzle guide plate has mirror symmetry about a first plane perpendicular to the fuel nozzle guide plate and mirror symmetry about a second plane parallel to the fuel nozzle guide plate. The fuel nozzle guide plate is attached to an air swirler, which is mounted in an opening of a bulkhead for mixing air and fuel at an upstream side of a combustor. | 01-13-2011 |
| 20110004388 | TURBOFAN TEMPERATURE CONTROL WITH VARIABLE AREA NOZZLE - A control system for a turbofan comprises a variable area nozzle for regulating core flow through the turbofan, an actuator, a temperature sensor, a flight controller and a nozzle control. The actuator is coupled to the variable area nozzle to regulate the core flow by positioning the variable area nozzle. The temperature sensor is positioned in the turbofan to sense a gas path temperature of the core flow. The flight controller is connected to the turbofan to make a thrust demand based on a flight condition of the turbofan. The nozzle control is connected to the flight controller and the actuator for directing the actuator based on the gas path temperature and the flight condition, such that the gas path temperature is controlled by adjusting the variable area nozzle to regulate the core flow while the turbofan meets the thrust demand. | 01-06-2011 |
| 20110000552 | DYNAMIC LEAK CONTROL FOR SYSTEM WITH WORKING FLUID - An organic rankine cycle system includes a sensor for sensing a condition indicative of pressure within the system and a control which responsively provides heat to said system when the pressure within the system is sensed to be at a predetermined threshold, near ambient pressure, during periods in which the system is shut down or preparing to operate. Provision is also made to remove the heat from the system when the pressure therein rises to a predetermined higher pressure threshold. | 01-06-2011 |
| 20110000086 | DEVICE AND METHOD FOR REPAIRING A FORWARD FRAME V-BLADE IN AN AIRCRAFT ENGINE - A repair device and method for repairing a forward frame v-blade on an aircraft engine are described in which the repair device replaces a section of the forward frame v-blade. In one embodiment, the repair device includes a supplemental part, which is secured to the forward frame v-blade. The supplemental part includes a central web and a pair of legs with a leg thickness that is greater than the thickness of corresponding flanges on the forward frame v-blade. Fasteners such as rivets are used to couple the supplemental part to the forward frame v-blade. | 01-06-2011 |
| 20100331140 | EPICYCLIC GEAR SYSTEM WITH LOAD SHARE REDUCTION - A journal bearing includes a main body and a connection portion. The main body extends radially outward from a central cavity to an interface surface. The connection portion supports an end of the main body and extends radially around the central cavity. The connection portion has a dimension selected as a proportion of a reference dimension of the main body to achieve a desired spring rate for the journal bearing. | 12-30-2010 |
| 20100331139 | EPICYCLIC GEAR SYSTEM WITH SUPERFINISHED JOURNAL BEARING - An epicyclic gear assembly includes a ring gear, a sun gear, at least one star gear enmeshed between the ring gear and sun gear, a carrier and a journal bearing. The carrier is disposed adjacent the rotatable sun gear and star gear. The journal bearing is disposed in the at least one star gear and connected to the carrier. The journal bearing has an outer radial surface with an amorphous surface finish of less than about 5 micro inches (127 micro mm) measured on an R | 12-30-2010 |
| 20100329835 | AIRFOIL WITH HYBRID DRILLED AND CUTBACK TRAILING EDGE - An apparatus for a gas turbine engine includes an airfoil, a metering opening for metering a cooling fluid, a cutback slot configured to deliver the cooling fluid from the metering opening, and a cooling hole. The airfoil defines a trailing edge, opposite first and second faces, and a mean camber line. The cutback slot is defined along the first face of the airfoil adjacent to the trailing edge and offset from the mean camber line of the airfoil. The cooling hole has an outlet that is located at the trailing edge and substantially aligned with the mean camber line of the airfoil. The cooling hole delivers a portion of the cooling fluid from the metering opening. | 12-30-2010 |
| 20100322782 | Nosecone bolt access and aerodynamic leakage baffle - A device includes a resilient member having a first end for blocking an access hole and a second end for attachment to a support surface, an opening extending through the resilient member and located between the first and second ends, and a bolt connection location positioned at the second end of the resilient member. The first end of the resilient member blocks the access hole and can be deflected to provide tool access to the bolt connection location through the first opening. | 12-23-2010 |
| 20100319358 | Nosecone ice protection system for a gas turbine engine - A system includes a first anti-icing electrothermal heater located at tip of a gas turbine engine nosecone, and a plurality of anti-icing electrothermal heater strips extending rearward from the first anti-icing electrothermal heater along the nosecone. | 12-23-2010 |
| 20100317478 | JOURNAL BEARING WITH SINGLE UNIT JUMPER TUBE AND FILTER - A journal bearing includes a central body having first passageway and a filter module. The central body extends axially and is adapted to be supported at each outer end. The first passageway extends generally axially through a portion of the central body. The filter module is disposed in the first passageway and is configured to trap debris from a lubricant fluid flowing therethrough. In one embodiment, the filter module is adapted to be inserted and removed from the first passageway as a single unit, thereby allowing the filter module to be connected and disconnected from fluid communication with a lubricant manifold positioned adjacent the journal bearing. | 12-16-2010 |
| 20100317477 | EPICYCLIC GEAR SYSTEM WITH IMPROVED LUBRICATION SYSTEM - A journal bearing that includes a central body having first and second passageways and a filter. The central body extends axially and is adapted to be supported at each outer end. The first passageway extends generally axially through a portion of the central body. The filter is disposed in the first passageway and is configured to trap debris from a lubricant fluid flowing therethrough. The second passageway extends between both the first passageway and an exterior surface of the central body to allow for delivery of lubricant fluid thereto. | 12-16-2010 |
| 20100310873 | PLASTIC WELDING USING FIBER REINFORCEMENT MATERIALS - A method of reinforcing a thermoplastic part includes softening a portion of the thermoplastic part to form a pool, embedding fibers in the softened pool, and re-solidifying the pool embedded with fibers into a weld that strengthens the thermoplastic part. | 12-09-2010 |
