Patent application number | Description | Published |
20130215260 | ACCURATE MACHINE TOOL INSPECTION OF TURBINE AIRFOIL - Inspecting features of an object including measuring features of the master part that represents a desired nominal dimensions to obtain a preliminary set of dimensional data. The preliminary set of data becomes a master data set. The same master part is measured on a second machine to obtain another set of data on the master part. The two data sets are compared and correction data is generated representing differences between the different data sets for the same master part measured on different machines. Subsequent parts are then measured on the second machine to obtain measurement data that is corrected based on the correction data. The application of the correction information provides for the use of 5-axis machines within a desired measurement capability. | 08-22-2013 |
20130223994 | OIL BAFFLE FOR GAS TURBINE FAN DRIVE GEAR SYSTEM - A turbine engine includes a housing supporting compressor, turbine sections, and an epicyclic gear train including a carrier. Sun gear and intermediate gears are arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubrication on the at least one of the sun gear and intermediate gears. | 08-29-2013 |
20130224471 | METHOD FOR FABRICATING A CERAMIC MATERIAL - A method for fabricating a ceramic material includes providing a mobilized filler material capable of infiltrating a porous ceramic matrix composite. The mobilized filler material includes at least one of a ceramic material and a free metal. The mobilized filler material is infiltrated into pores of the porous ceramic matrix composite. The mobilized filler material is then immobilized within the pores of the porous ceramic matrix composite. | 08-29-2013 |
20130276424 | Low Noise Compressor Rotor for Geared Turbofan Engine - A gas turbine engine has a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion. The low pressure turbine portion drives the low pressure compressor portion and the fan. A gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor portion. At least one of the low pressure turbine portion and low pressure compressor portion has a number of blades in each of a plurality of rows. The blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the at least one of the low pressure turbine portion and/or the low pressure compressor sections: (number of blades×rotational speed)/60≧5500. The rotational speed is an approach speed in revolutions per minute. | 10-24-2013 |
20130298565 | GAS TURBINE ENGINE SYSTEMS AND RELATED METHODS INVOLVING MULTIPLE GAS TURBINE CORES - Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly. | 11-14-2013 |
20130323417 | 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. | 12-05-2013 |
20140056685 | ANNULAR TURBOMACHINE SEAL AND HEAT SHIELD - An example annular turbomachine seal includes an attachment portion, a first leg radially inward of the attachment portion, and a second leg radially inward of the attachment portion and configured to block flow from directly contacting an attachment associated with a bearing compartment of a turbomachine. | 02-27-2014 |
20140079546 | GAS TURBINE ENGINE FRONT CENTER BODY ARCHITECTURE - Disclosed is a gas turbine engine configured to include a propulsor section having a fan drive geared architecture. The gas turbine engine includes a gas generator section that includes a compressor section having a first rotor rotationally mounted to a first spool. The engine further includes a bearing package adapted for rotationally supporting the first spool. The bearing package is configured for supporting the first rotor during at least one of (1) a period prior to installation of the fan drive geared architecture into the engine and (2) a period after removal of the fan drive geared architecture from the engine. | 03-20-2014 |
20140082940 | METHOD AND FIXTURE FOR AIRFOIL ARRAY ASSEMBLY - An example method of assembling a turbomachine airfoil array includes, among other things, securing a partial airfoil array within a fixture, the partial airfoil array having at least one existing airfoil extending radially between an inner and an outer fairing and an open area where at least one existing airfoil has been removed. The method includes mounting a positioning saddle relative to a base of the fixture, the positioning saddle aligned with the open area, holding a replacement airfoil using the positioning saddle, applying a curable material at an interface between the replacement airfoil and the inner and outer fairing, and curing the curable material while maintaining a relative position between the replacement airfoil and the inner and outer fairing. | 03-27-2014 |
20140083076 | Gas Turbine Engine With Geared Turbofan and Oil Thermal Management System - A lubricant supply system for a gas turbine engine has a lubricant lube pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear for an associated gas turbine engine. A method and apparatus are disclosed. | 03-27-2014 |
20140083079 | GEARED TURBOFAN PRIMARY AND SECONDARY NOZZLE INTEGRATION GEOMETRY - A disclosed example geared turbofan engine includes a fan section including a plurality of fan blades rotatable about an axis and a core engine section defined about an engine axis. The core engine section includes a primary nozzle including a primary outer diameter at a primary nozzle trailing edge and a primary maximum inner diameter forward of the primary trailing edge. A bypass passage is defined between an inner nacelle surrounding the core engine section and an outer nacelle and includes a secondary nozzle. The secondary nozzle includes an outer diameter at a secondary nozzle trailing edge and a secondary maximum inner diameter forward of the secondary trailing edge. A ratio between the maximum inner diameter of the primary nozzle and an outer diameter at the trailing edge of the primary nozzle and a ratio between the maximum inner diameter of the secondary trailing edge and the outer diameter at the trailing edge of the secondary nozzle are both less than about 0.700. | 03-27-2014 |
20140083114 | TURBINE BLADE ROOT PROFILE - A turbine blade for a gas turbine engine includes an airfoil that extends in a first radial direction from a platform. A root extends from the platform in a second radial direction and has opposing lateral sides that provide a firtree-shaped contour. The contour includes first, second and third lobes on each of the lateral sides and that tapers relative to the radial direction away from the platform. The first, second and third lobes each provide contact surfaces arranged at about 45° relative to the radial direction. A contact plane on each lateral side at an angle of about 11° relative to the radial direction defining a contact point on each of the contact surfaces. The first, second and third lobes each include first, second and third grooves that are substantially aligned with one another along an offset plane spaced a uniform offset distance from the contact plane. | 03-27-2014 |
20140086741 | TURBOMACHINE HYBRID LIFT-OFF FACE SEAL - An example of a turbomachine seal includes, among other things, a sealing member configured to be influenced by both hydrostatic and hydrodynamic forces when providing a seal. | 03-27-2014 |
20140090388 | OFF-TAKE POWER RATIO - An example method of allocating power within a gas turbine engine includes driving an off-take power delivery assembly using a first amount of power from a spool, the first amount of power corresponding to an off-take power requirement of a gas turbine engine; and driving the spool of the gas turbine engine using a second amount of power, wherein a ratio of the first amount of power to the second amount of power is greater than or equal to 0.009. | 04-03-2014 |
20140090392 | MODEL BASED FUEL-AIR RATIO CONTROL - A gas turbine engine comprises a compressor, a combustor, a turbine, and an electronic engine control system. The compressor, combustor, and turbine are arranged in flow series. The electronic engine control system is configured to estimate combustor fuel-air ratio based on a realtime model-based estimate of combustor airflow, and commands engine actuators to correct for a difference between the estimated combustor fuel-air ratio and a limit fuel-air ratio selected to avoid lean blowout. | 04-03-2014 |
20140093363 | LINER AND METHOD OF ASSEMBLY - An assembly includes a plurality of vanes, a forward liner segment, and an aft liner segment. The forward liner segment and the aft liner segment are mounted to the plurality of vanes and each segment comprises an arc of less than 360° in length. | 04-03-2014 |
20140109546 | GROUNDING FOR FAN BLADES ON AN UNDERBLADE SPACER - A fan rotor includes a rotor body with at least one slot receiving a fan blade. The fan blade has an outer surface, at least at some areas, formed of a first material and an airfoil extending from a dovetail. The dovetail is received in the slot. A spacer is positioned radially inwardly of the dovetail biasing the fan blade against the slot. The spacer includes a grounding element, which is in contact with a portion of the dovetail formed of a second material that is more electrically conductive than the first material. The grounding element is in contact with a rotating element that rotates with the rotor. The rotating element is formed of a third material. The first material is less electrically conductive than the third material. The grounding and rotating elements form a ground path from the portion of the dovetail into the rotor. | 04-24-2014 |
20140109548 | HIGH PRESSURE ROTOR DISK - A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships. | 04-24-2014 |
20140112786 | Reversible Blade Damper - A damper body extends between an axial first end, an opposing axial second end, a first lateral side, and an opposing second lateral side. A first pair of lugs are disposed at the first axial end and a second pair of lugs are disposed at the second axial end. In the first lug pair, a lug has an axially outermost foot which extends laterally beyond a lateral wall. The second lug pair has a lug with an axial outermost point with a foot extending laterally beyond a lateral wall. The foot on the first axial end and the foot on the second axial end are on opposed lateral sides. | 04-24-2014 |
20140119887 | FLUID-COOLED SEAL ARRANGEMENT FOR A GAS TURBINE ENGINE - A rotational component for a gas turbine engine includes at least one passage with a semi-spherical end. | 05-01-2014 |
20140127007 | GAS TURBINE ENGINE ROTOR SEAL - A rotary seal for sealing a bladed rotor of a gas turbine engine to a stator thereof comprises a sealing element and a sealing element support comprising a radially outer edge portion on which the sealing element is fixed, a radially inner mounting portion adapted to be attached to a supported rotor hub, and a flexible, medial web portion extending between the radially outer edge portion and the radially inner mounting portion. | 05-08-2014 |
20140127013 | GAS TURBINE ENGINE AIRFOIL COOLING CIRCUIT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that extends between a leading edge, a trailing edge, a pressure side wall and a suction side wall. A cooling circuit is disposed inside of the airfoil. The cooling circuit includes a first core cavity that radially extends inside of the airfoil. A first axial skin core is in fluid communication with the first core cavity at a first location of the first axial skin core and a second core cavity is in fluid communication with the first axial skin core at a second location of the first axial skin core. | 05-08-2014 |
20140133973 | POWER CABLE AND CONNECTOR ARRANGEMENT FOR A GAS TURBINE ENGINE - A gas turbine engine component assembly according to an exemplary aspect of the present disclosure includes, among other things, a nacelle including a moveable portion. A generator is configured to be received within a space established by the nacelle. A plurality of cables are at least partially supported on the moveable portion of the nacelle, the plurality of cables being selectively electrically coupled with the generator. | 05-15-2014 |
20140137538 | Fast Response Bypass Engine - A gas turbine engine has a compressor, a fan for delivering air into the compressor and into a bypass duct, a combustion section and a turbine section. A control for the gas turbine engine is programmed to change a fueling level and position at least one effector that may be moved to unique positions in a coordinated fashion upon receipt of a command to change thrust. The thrust provided by the engine is changed without a reduction in an airflow stability margin compared to a thrust change commanded only by a fueling change. Some aspects of the positioning are transitory. | 05-22-2014 |
20140140818 | GAS TURBINE ENGINE FRONT CENTER BODY ARCHITECTURE - A method for servicing a gas turbine engine includes disassembling a bearing compartment, providing access from a forward side of the gas turbine engine to a gearbox contained within said bearing compartment. The gas turbine engine provides a core flow path that extends from the forward side aftward in a core flow direction. The method includes servicing a component located within the bearing compartment. A method for servicing a gas turbine engine includes providing access from a forward side of a front center body assembly, and servicing a component located within a bearing compartment aft of the front center body assembly. The gas turbine engine includes a seal package located aft of a bearing package. A front wall is mounted to the front center body support and is removable from a front center body support to access at least one of the gearbox, the bearing package and the seal package. | 05-22-2014 |
20140140823 | HYBRID GAS TURBINE BEARING SUPPORT - A hybrid radial gas turbine engine component comprises an inner hub portion joined to an outer ring portion. The inner hub portion is a first alloy and operates at temperatures less than 1200° F. The outer ring portion is a second alloy and is designed to withstand extended periods at temperatures greater than 1200° F. | 05-22-2014 |
20140140824 | OIL SYSTEM BEARING COMPARTMENT ARCHITECTURE FOR GAS TURBINE ENGINE - A gas turbine engine with a geared architecture includes a multiple of bearing compartments and at least one carbon seal that seals at least one side of each of the multiple of bearing compartments. | 05-22-2014 |
20140140827 | Full Ring Inner Air-Seal With Locking Nut - A vane assembly has a vane including an inner platform with a mount rail extending radially inwardly toward an engine center axis. An air seal is attached to the inner platform. The air seal has a single-piece ring extending circumferentially about the engine center axis. A ring nut secures the air seal to the inner platform. A biasing member cooperates with the ring nut and air seal to seal fore and aft locations on the vane. | 05-22-2014 |
20140140828 | BLAST SHIELD FOR HIGH PRESSURE COMPRESSOR - A high pressure compressor includes an inner case wall, a plurality of flanges extending outwardly from the inner case wall, a shield that surrounds at least a portion of at least one of the plurality of flanges to define a chamber that at least partially encases at least one of the plurality of flanges, and a passage at least partially defined by the shield that allows air to circulate into the chamber. | 05-22-2014 |
20140142889 | Throat Area Calculation for a Section of a Gas Turbine Engine - A method of calculating a throat area of a section of a machinery, according to an exemplary aspect of the present disclosure includes, among other things, outlining a boundary of the throat area of the section, selecting a plurality of inspection points along the boundary of the throat area, dividing the throat area into a plurality of sections, calculating an individual area of each of the plurality of sections and summing the individual areas of each of the plurality of sections to calculate the throat area. | 05-22-2014 |
20140147249 | GAS TURBINE ENGINE ROTOR DRAIN FEATURE - An assembly for a gas turbine engine includes a rotor formed by a drum and a hub. The drum has a generally conically shaped inner surface that defines a flow path along a first cavity. The hub is connected to the drum and has a conically shaped inner surface that defines a flow path along a second cavity. | 05-29-2014 |
20140147263 | TURBINE VANE WITH MISTAKE REDUCTION FEATURE - A turbine vane for a gas turbine engine has an inner platform, an outer platform, at least one airfoil extending between the inner and outer platforms, and a tab radially extending inward from a front side of the inner platform. The tab contains a mounting aperture and an identification aperture that identifies an engine in which the turbine vane may be installed. | 05-29-2014 |
20140147271 | Pressure Seal With Non-Metallic Wear Surfaces - A gas turbine engine has a pair of components having a high-pressure chamber on one side, and a low pressure chamber on an opposed side. A three-sided seal has one side in sealing contact with each of said components. A third side is associated with a third component. At least one non-metallic wear surface is positioned between one of the three sides of the seal and its respective component. | 05-29-2014 |
20140147287 | TRAILING EDGE AND TIP COOLING - A gas turbine engine component includes an airfoil extending radially from a root section to a tip section and having a trailing edge cooling passageway and first, second and third flow dividers in the cooling passageway. The first, second and third flow dividers have longitudinal axes that are angled based upon a position of the flow divider relative to the tip section of the airfoil. | 05-29-2014 |
20140154068 | Endwall Controuring - An airfoil array is disclosed. The airfoil array may include an endwall, and a plurality of airfoils radially projecting from the endwall. Each airfoil may have a first side and an opposite second side extending axially in chord between a leading edge and a trailing edge. The airfoils may be circumferentially spaced apart on the endwall thereby defining a plurality of flow passages between adjacent airfoils. The airfoil array may further include a convex profiled region extending from the endwall adjacent to the first side of at least one of said plurality of airfoils near the leading edge of the at least one of said plurality of airfoils, and a concave profiled region in the endwall and extending across said at least one of said plurality of flow passages. | 06-05-2014 |
20140161593 | TURBINE ENGINE ALIGNMENT ASSEMBLY - An augmentor assembly for a gas turbine engine includes an alignment assembly disposed along an axis having a base and a plate. The base includes a plurality of grooves. The plate includes a plurality of flanges in mating engagement to the plurality of grooves. A first component and a second component are attached to the alignment assembly. The mating engagement of the plurality of grooves and the plurality of flanges is arranged to provide incremental adjustment to define a position of the first component relative to the second component. | 06-12-2014 |
20140161616 | MULTI-PIECE BLADE FOR GAS TURBINE ENGINE - A blade assembly for a gas turbine engine includes a rim seal including leading and trailing edge seal portions joined to one another by an axial portion. The leading and trailing edge seal portions and the axial portion together provide a notch. A blade has a root received in the notch. A rotating stage of a gas turbine engine includes a rotor including a slot. A rim seal includes leading and trailing edge seal portions adjoined to one another by an axial portion and providing a notch. A blade has a root received in the notch. A method of assembling a rotor stage includes inserting a root of a blade into a notch of a rim seal, and sliding the rim seal and blade into a rotor slot. | 06-12-2014 |
20140165530 | Bi-Directional Auxiliary Lubrication System - A bi-directional auxiliary lubrication system which allows lubricant to be supplied to moving engine components after a loss of lubricant pressure from a main lubricant tank is disclosed. In a gas turbine engine, the lubrication system may siphon compressed air from a compressor to draw lubricant from a reserve lubricant tank and deliver that lubricant to the engine components. The same conduits used by the lubrication during normal operations are utilized in an opposite direction to provide the flow of lubricant from the reserve lubricant tank during such auxiliary or low-lubricant-pressure operations. | 06-19-2014 |
20140165570 | OSCILLATING HEAT PIPE FOR THERMAL MANAGEMENT OF GAS TURBINE ENGINES - An oscillating heat pipe of a gas turbine engine includes a plurality of channels that define a continuous loop through which a fluid flows, and an evaporator of a gas turbine engine. The fluid flows through the evaporator to accept heat from a first fluid. The first fluid is located near or in an engine core. The oscillating heat pipe also includes condenser of the gas turbine engine. The fluid flows through the condenser to reject heat to a second fluid, and the second fluid is located outwardly of the engine core. | 06-19-2014 |
20140165586 | TURBINE START METHOD - A method of starting an engine includes the steps of starting a fuel supply pump, opening a fuel solenoid to allow fuel flow to a combustor to at least partially begin, and beginning to operate a starter motor to drive a shaft associated with the engine. Then an ignitor is excited to spark in the combustor. An engine is also disclosed. | 06-19-2014 |
20140166255 | Closure of Cooling Holes with a Filing Agent - A method for filling cooling holes in a component of a gas turbine engine is disclosed. The component may include a plurality of first cooling holes penetrating the wall of the component. The method may comprise the steps of exposing the outer surface of the component, filling the plurality of first cooling holes of the component with a filling agent, curing the filling agent to block the passage of air through the cooling holes, and applying a thermal barrier coating over the surface of the component. The method may further include installing a second plurality of cooling holes, the second plurality of cooling holes penetrating the thermal barrier coating and the wall of the component and allow air to pass therethrough. | 06-19-2014 |
20140169934 | ANTI-ICE SUPPLY SYSTEM FOR INLET GUIDE VANES - Disclosed is a gas turbine engine including a plurality of inlet guide vanes. The gas turbine engine further includes an inlet case supporting each of the plurality of inlet guide vanes, and the inlet case is provided with at least one passageway in communication with a source of fluid. The at least one passageway is configured to communicate the fluid to each of the plurality of inlet guide vanes. | 06-19-2014 |
20140169941 | FORWARD COMPARTMENT BAFFLE ARRANGEMENT FOR A GEARED TURBOFAN ENGINE - A baffle within a gas turbine engine includes a cylindrical flange with a multiple of holes at least partially therearound. | 06-19-2014 |
20140169976 | ROTOR BLADE ROOT SPACER FOR ARRANGING BETWEEN A ROTOR DISK AND A ROOT OF A ROTOR BLADE - A rotor assembly is provided that includes a rotor blade, a root spacer, and a rotor disk with a slot. The rotor blade includes a blade root arranged within the slot. The blade root includes a root base segment and a pair of root side segments. The root base segment is laterally separated from the rotor disk by the root side segments. The root spacer is arranged within the slot, and includes a side surface that extends radially between an inner surface and an outer surface. The side surface is approximately laterally aligned with an intersection between the root base segment and a first of the root side segments. The outer surface engages the root base segment. | 06-19-2014 |
20140178196 | TURBOMACHINE WITH CLAMP COUPLING SHAFT AND ROTOR HUB TOGETHER - A turbomachine includes a rotor hub that has a central opening there through. A shaft extends through the central opening. A clamp is coupled with the shaft and the rotor hub such that the rotor hub is rotatable with the shaft. | 06-26-2014 |
20140186165 | ADHESIVE PATTERN FOR FAN CASE CONFORMABLE LINER - A section of a gas turbine engine includes a case structure and a liner. The liner is attached to the case structure by a compliant adhesive. The adhesive covers at least a portion of the liner surface area, and can have interruptions used to tune the compliance of the adhesive layer. | 07-03-2014 |
20140196433 | GAS TURBINE ENGINE COMPONENT PLATFORM COOLING - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least one of the leading edge portion and the trailing edge portion of the outer surface of the platform. | 07-17-2014 |
20140196434 | Shroud for Sealing Conduit Feed-Through - A system has a surface intended to separate two chambers within the system. The surface has an aperture for allowing passage of at least one communication conduit. A shroud is positioned on the surface at the aperture, and has at least two portions defining a central opening to allow the communication conduit to pass through the aperture and shroud. The two portions of shroud have mating clamp ears in contact with each other. Securement members tighten the clamp ears against each other to provide a seal at an end of the shroud remote from the surface. | 07-17-2014 |
20140196471 | FAN DRIVE GEAR SYSTEM FLEXIBLE SUPPORT FEATURES - A disclosed gear assembly support for a gas turbine engine includes a first portion configured for attachment to a case of the gas turbine engine and a second portion configured for supporting a gear assembly. The support further includes a snap portion defining a fit within the case. The snap portion includes a tunable feature for adjusting a fit within the case. A torque reacting portion of the support transfers torque from the second portion to the first portion separate from the snap portion and include separately tunable features fore adjusting the snap fit independent of the torque transfer portions. | 07-17-2014 |
20140196472 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 07-17-2014 |
20140196473 | GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE - A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine. | 07-17-2014 |
20140199177 | AIRFOIL AND METHOD OF MAKING - An airfoil includes leading and trailing edges, a first exterior wall extending from the leading edge to the trailing edge and having inner and outer surfaces, a second exterior wall extending from the leading edge to the trailing edge generally opposite the first exterior wall and having inner and outer surfaces, and cavities within the airfoil. A first cavity extends along the inner surface of the first exterior wall and a first inner wall and has an upstream end and a downstream end, and a feed cavity is located between the first inner wall and the second exterior wall. | 07-17-2014 |
20140202170 | Composite Articles and Methods - An article has a polymeric substrate and a coating system. The coating system includes a metallic plating and a polymeric coating atop the metallic plating. The metallic plating has a thickness of at least 0.05 mm. | 07-24-2014 |
20140208714 | Thermal Management for Gas Turbine Engine - A gas turbine engine has an inlet duct, which is configured to communicate with an inlet to a compressor. The inlet duct is further configured to communicate air outwardly of an outer casing of the gas turbine engine, and to pass the air along an axial length of the gas turbine engine to cool a component associated with the gas turbine engine. | 07-31-2014 |
20140208771 | GAS TURBINE ENGINE COMPONENT COOLING ARRANGEMENT - A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion having an exterior surface and an internal surface. A cavity is disposed inside of the body portion. A cooling hole extends between the exterior surface and the internal surface and includes a metering section having an outlet and an inlet. The inlet is shaped dissimilar to the outlet. | 07-31-2014 |
20140208837 | GAS TURBINE ENGINE VANE EMBEDDED BEAM INTERRUPT OPTICAL TIP-TIMING PROBE SYSTEM - A method of observing an airflow passage within a gas turbine engine includes locating a support in view of an airflow passage and housing an optical fiber within the support. | 07-31-2014 |
20140212260 | Airfoil Assembly with Paired Endwall Contouring - An airfoil assembly is disclosed. The airfoil assembly may comprise a series of airfoils such as vanes radially extending between an ID endwall and a substantially concentric OD endwall. The airfoils are circumferentially spaced apart and define a plurality of fluid flow passages between adjacent airfoils. The ID endwall and the OD endwall define contoured regions to optimize fluid flow through the passages. A geometric relationship between the ID endwall contoured regions and the OD endwall contoured regions assures that the throat area generally increases or decreases in the direction of fluid flow. | 07-31-2014 |
20140212261 | LIGHTWEIGHT SHROUDED FAN - A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan with a plurality of fan blades rotatable about an axis and a shroud and a speed change device in communication with the fan. | 07-31-2014 |
20140212262 | VARIABLE OUTER AIR SEAL SUPPORT - A variable outer air seal support system according to an exemplary aspect of the present disclosure includes, among other things, a case having a plurality of slots, and an extension of a variable outer air seal segment. The extension provides at least one extension aperture. A connector pin is configured to move within the slot to move the variable outer air seal segment from a first position to a second position. The variable outer air seal segment overlaps a circumferentially adjacent variable outer air seal segment more in the first position than in the second position. | 07-31-2014 |
20140212276 | VARIABLE OUTER AIR SEAL FLUID CONTROL - An example turbomachine system includes a first variable outer air seal including at least one channel. The first variable outer air seal configured to selectively communicate a fluid in response to movement of a second variable outer air seal relative to the first variable outer air seal. An example fluid control method includes selectively covering a channel inlet using a variable outer air seal to control flow through the channel. | 07-31-2014 |
20140212281 | Flow Feed Diffuser - A feed diffuser is disclosed. The diffuser may comprise a base disposed orthogonal to a flow, a body extending from the base toward the flow, and an apex provided at a termination of the body in a direction toward the flow. The body may decrease in cross-sectional area from the base to the apex. | 07-31-2014 |
20140212292 | Airfoil Member and Composite Platform Having Contoured Endwall - An airfoil member and platform assembly for a gas turbine engine is provided. The platform to which the airfoil member is anchored is made of a composite material and includes an endwall defining a contoured region for improved aerodynamics. The contoured region influences the flow of gases through the flow passages between the airfoil members, thereby reducing endwall losses due to horseshoe vortexing. The composite material may be woven ceramic matrix composite fibers infiltrated with a ceramic matrix material or woven organic matrix composite fibers infiltrated with an organic matrix material. | 07-31-2014 |
20140212297 | GAS TURBINE ENGINE SERPENTINE COOLING PASSAGE WITH CHEVRONS - A gas turbine engine component includes a structure having a cooling passage providing upstream and downstream portions separated from one another by an inner wall and fluidly connected by a bend. First and second trip strips are respectively arranged in the upstream and downstream portions. The first trip strips are arranged at a first spacing from one another. The second trip strips are arranged at a second spacing from one another. A turbulence promoter is arranged in the bend and at a third spacing from the first trip strips that is different than the first spacing. The turbulence promoter is arranged at a fourth spacing from the second trip strips that is different than the second spacing. | 07-31-2014 |
20140212298 | GAS TURBINE ENGINE TURBINE BLADE TIP COOLING - An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A camber line at the tip extends from the leading edge to the trailing edge. Pressure and suction side shelves are arranged in the exterior surface on opposing sides of the camber line respectively in the pressure and suction side walls. A plateau is proud of and separates the pressure and suction side shelves. The plateau is arranged along the camber line and extends to the leading edge. | 07-31-2014 |
20140216005 | Gas Turbine Engine Assembly Including A Thrust Reverser - An exemplary gas turbine engine assembly includes a thrust reverser that is selectively moveable between a stowed position and a thrust reversing position. The thrust reverser includes an outer surface having a first outer surface area when the thrust reverser is in the stowed position and a second, smaller outer surface area when the thrust reverser is in the thrust reversing position. | 08-07-2014 |
20140216044 | GAS TURBINE ENGINE COMBUSTOR HEAT SHIELD WITH INCREASED FILM COOLING EFFECTIVENESS - A heat shield for a gas turbine engine includes a hot side with one or more raised features that extend therefrom. | 08-07-2014 |
20140219778 | GAS TURBINE ENGINE AIRFOIL WITH VANE PLATFORM COOLING PASSAGE - A stator vane for a gas turbine engine includes an airfoil extending in a radial direction and supported by a platform having a gas flowpath surface. A cooling passage is arranged in the platform and includes a circumferential passage that is fluidly connected to an inlet passage extending through and edge of the platform, and film cooling holes extending from the gas flowpath surface to the circumferential passage, radial extending passage through the edge of the platform. A void is interconnected to at least one of the radially extending passage and the inlet passage. | 08-07-2014 |
20140219792 | PYLON MATCHED FAN EXIT GUIDE VANE FOR NOISE REDUCTION IN A GEARED TURBOFAN ENGINE - A disclosed fan section of a gas turbine engine includes a fan rotor having a plurality of fan blades and a duct defining a passageway aft of the fan rotor. A fan exit guide vane is disposed within the duct downstream of the fan blades. The fan exit guide vane includes a plurality of exit guide vanes positioned downstream of the fan rotor with at least two of the plurality of exit guide vanes including different aft geometries for guiding airflow through the passage to reduce pressure distortions at the fan blades. | 08-07-2014 |
20140219793 | HEALTH MONITORING FOR HOLLOW BLADES - A system for detecting a defect in a hollow blade includes a pressure sensor mounted within a cavity of the hollow blade, wherein the pressure sensor measures a pressure within the cavity and provides an output indicative of measured pressure; and a controller to receive the output from the pressure sensor, wherein the controller indicates a defect in the hollow blade based upon the received output. | 08-07-2014 |
20140219794 | BORE CAVITY THERMAL CONDITIONING SYSTEM - A gas turbine engine includes a rotor having at least one rotor bore cavity and a heat exchanger. The heat exchanger is fluidically connected to a conditioning air source and to the rotor bore cavity for flowing air from the conditioning air source to the rotor bore cavity. The heat exchanger is also fluidically connected to a hot fluid source and a cool fluid source for selectively heating and cooling conditioning air flowing to the rotor bore cavity. | 08-07-2014 |
20140219808 | SHEATH WITH EXTENDED WINGS - A sheath for a fan airfoil having a leading edge and a trailing edge in a chordwise direction, a tip and a root in a spanwise direction, a suction side and a pressure side includes a solid portion to wrap around the airfoil leading edge; a first wing attached to the suction side of the airfoil; and a second wing attached to the pressure side of the airfoil. At least one of the first wing and the second wing extends at least about 35% of the chord of the airfoil. | 08-07-2014 |
20140219815 | MULTI-LOBED COOLING HOLE - A gas turbine engine component subjected to a flow of high temperature gas includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe, a second lobe and a transition region. The first lobe diverges longitudinally and laterally from the metering section and has a first shape. The second lobe is generally opposite the first lobe and diverges longitudinally and laterally from the metering section and has a second shape different from the first shape. The transition region is positioned between the first and second lobes and includes a downstream end adjacent the outlet. | 08-07-2014 |
20140237990 | VARIABLE AREA FAN NOZZLE WITH WALL THICKNESS DISTRIBUTION - A gas turbine engine includes a core engine that has at least a compressor section, a combustor section and a turbine section disposed along a central axis. A fan is coupled to be driven by the turbine section. A fan nacelle is arranged around the fan, and a bypass passage extends between the fan nacelle and the core engine. A variable area fan nozzle (VAFN) extends at least partially around the central axis and defines an exit area of the bypass passage. The VAFN is selectively movable to vary the exit area. The VAFN includes a body that defines an airfoil cross-section shape. The VAFN includes a wall that has a mechanical property distribution in accordance with a computer-simulated vibration profile of a flutter characteristic of the VAFN. | 08-28-2014 |
20140238027 | THERMALLY COMPLIANT DUAL WALL LINER FOR A GAS TURBINE ENGINE - A stiffener for a liner assembly of a gas turbine engine includes a second resilient member arranged on a first resilient member at substantially ninety (90) degree orientation. | 08-28-2014 |
20140241857 | FLOW DIVERTER TO REDIRECT SECONDARY FLOW - An assembly for a gas turbine engine includes a seal and a flow diverter. The flow diverter is disposed adjacent the seal to direct a secondary gas flow that passes across the seal away from a rotor cavity such that the secondary gas flow travels back toward a main gas flow path of the gas turbine engine. | 08-28-2014 |
20140241874 | WEAR LINER SPRING SEAL - A wear liner for a stator vane includes a shaped member, comprising a U-shaped portion that receives a foot of a stator vane. A first end of the wear liner is biased into contact with the stator vane. A spring seal extends from the U-shaped portion and into bias contact with an inner surface of the support structure to further prevent leakage between the stator foot and case support structure. | 08-28-2014 |
20140241897 | ALUMINUM BRAZING OF HOLLOW TITANIUM FAN BLADES - A fan blade includes first and second titanium portions that are secured to one another with an aluminum alloy braze. A method of manufacturing a fan blade includes providing first and second titanium portions, applying an aluminum alloy braze to at least one of the first and second titanium portions, and heating the fan blade to melt the aluminum alloy braze and join the first and second portions to one another to provide a fan blade with an airfoil exterior contour. | 08-28-2014 |
20140245749 | Nacelle Anti-Ice Valve Utilized as Compressor Stability Bleed Valve During Starting - A gas turbine engine has a compressor section received within an inner housing. An is an outer housing is spaced radially outwardly of the inner core housing. A nacelle has an anti-icing system which taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve is opened at startup of the gas turbine engine to assist compressor stability. | 09-04-2014 |
20140245750 | CIRCUMFERENTIALLY RETAINED FAIRING - An assembly for a gas turbine engine includes a frame, a mount, and a fairing. The mount is attached to the frame and the fairing is connected to the mount. The fairing and mount have mating anti-deflection features that engage to prevent circumferential movement of the fairing relative to the frame. | 09-04-2014 |
20140245751 | PASSAGES TO FACILITATE A SECONDARY FLOW BETWEEN COMPONENTS - An annular mount for a gas turbine engine includes a first flange and a second flange. The first flange has an arcuate shape, and the second flange extends from the first flange. The second flange has an interface surface along at least one side and has a plurality of mounting apertures extending therethrough. The channels extend along a length of the interface surface from a first edge to a second edge. | 09-04-2014 |
20140245753 | GAS TURBINE ENGINE ROTOR BLADE - A rotor blade for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil extending in span between a root region and a tip region and a tip portion extending at an angle from the tip region of the airfoil. | 09-04-2014 |
20140246514 | TURBOMACHINE THRUST REVERSER - An example thrust reverser of a gas turbine engine is configured to connect to an aircraft wing via a pylon via one or more thrust reverser mounts located adjacent to a top circumferential apex of the engine according to an exemplary aspect of the present disclosure includes, among other things, a first cowl moveable between a stowed position and a deployed position relative to a second cowl. The first cowl in the deployed position configured to permit thrust to be redirected from an engine to slow the engine. The first cowl forming a portion of a substantially annular encasement of the engine. The first cowl directly interfaces with second cowl of the encasement at a cowl interface position that is more than 18 degrees circumferentially offset from the top circumferential apex when the first cowl is in the stowed position. | 09-04-2014 |
20140248115 | Active Clearance Control System with Zone Controls - An ACC system and method of using such for changing a turbine blade to BOAS gap on an aircraft engine is disclosed. The ACC system may comprise a first ring, a first supply line and a first flow control assembly. The first ring may be configured to substantially encircle a portion of a case assembly that is disposed around an aircraft engine turbine. The first ring may include a plurality of segments that each define a chamber, an inlet port and a plurality of outlet ports. At least a portion of the outlet ports may be configured to be disposed adjacent to the case. The first supply line may be operatively connected to a first segment of the plurality of segments. The first flow control assembly may be operatively connected to the first supply line and configured to meter the flow of cool air into the first segment. | 09-04-2014 |
20140248116 | ACOUSTIC TREATMENT IN AN UNDUCTED AREA OF A GEARED TURBOMACHINE - An exemplary geared turbomachine assembly includes an acoustic treatment that is in an unducted area of a geared turbomachine. | 09-04-2014 |
20140248126 | INTER-MODULE FINGER SEAL - An assembly for a gas turbine engine includes a first module, a second module, and a finger seal. The second module is connected to the first module along a joint. The finger seal is connected to the first module at a fixed end and has a free end that cantilevers from the fixed end to contact the second module. The finger seal acts to seal the joint between the first module and second module from an ingestion gas flow of the gas turbine engine. | 09-04-2014 |
20140248127 | TURBINE ENGINE COMPONENT WITH DUAL PURPOSE RIB - An assembly for a gas turbine engine includes a component, a fairing, and a seal. The fairing is disposed adjacent to the component and defines a primary gas flow path. The fairing has a rib that is located outside of the primary flow path and extends from an outer surface of the fairing. The seal is disposed between the rib and the component. | 09-04-2014 |
20140248128 | SEALS FOR A CIRCUMFERENTIAL STOP RING IN A TURBINE EXHAUST CASE - A turbine seal system comprises an annular structural frame, a circumferential ring, a fairing and a seal. The circumferential ring is joined to the annular structural frame. The fairing is disposed within the annular structural frame and is engaged with the circumferential ring to limit circumferential rotation of the fairing with respect to the annular structural frame. The seal extends between the fairing and the circumferential ring. In one embodiment, the structural component comprises a ring-strut-ring turbine exhaust case. | 09-04-2014 |
20140248130 | Seal Hook Mount Structure With Overlapped Coating - A seal is positioned radially outwardly of a compressor blade. The seal has a seal hook facing in a downstream direction. A seal mount has a mount hook for receiving the seal hook in a gap. The hook face in an upstream direction. The seal mount has an upstream portion extending from an upstream end into the gap. The mount hook extends into a web which extends in a downstream direction. A thermal barrier coating on the upstream portion of the seal mount extends to a downstream end. There is a thermal barrier coating on a radially inner surface of the mount hook, and along the web to a downstream coating end. An upstream end of the coating on the inner surface of the hook being formed axially upstream of the downstream end of the coating on the upstream portion. | 09-04-2014 |
20140248152 | PLATE FOR DIRECTING FLOW AND FILM COOLING OF COMPONENTS - An assembly includes a gas turbine engine component and a plate. The plate is spaced from a surface of the component and generally conforms to the shape of the surface. The plate and component form a passageway that allows for passage of a secondary gas flow between the component and the plate. | 09-04-2014 |
20140248156 | Composite Articles and Methods - An article comprises: a substrate having a matrix and a plurality of metallic members partially embedded in the matrix; and a metallic layer bonded to exposed portions of the metallic members. | 09-04-2014 |
20140255144 | FLANGED BUSHING FOR VARIABLE VANE - An example turbomachine bushing a bushing having a wear surface configured to interface directly with a variable vane assembly to limit radially inward movement of the variable vane assembly, wherein the variable vane assembly is moveable axially between a first position contacting the wear surface and a second position spaced from the wear surface. | 09-11-2014 |
20140255158 | COATING PROCESS FOR GAS TURBINE ENGINE COMPONENT WITH COOLING HOLES - A method of coating a component having a multiple of cooling holes includes removing at least a portion of a prior coating; directing a gas through at least one of the multiple of cooling holes; and applying a coat layer while directing the gas through at least one of the | 09-11-2014 |
20140255162 | Adjustable Floating Oil Channel for Gas Turbine Engine Gear Drive - A turbine is operably connected to drive a compressor, and to drive a fan through a gear drive. A number of intermediate gears connecting an output shaft of the turbine to a fan drive shaft for the fan. An oil channel collects oil thrown outwardly of the gear drive. A bearing support mounts bearings supporting the fan drive shaft. The oil channel and the bearing support each include mating faces that are bolted together by a plurality of bolts. The bolts extend through oil channel holes in the mating face of the oil channel. The oil channel holes have one dimension which closely receives the bolts and another dimension which is larger than an outer diameter of the extending portion of the bolts, such that the bolts may adjust radially within the oil channel holes. | 09-11-2014 |
20140255168 | INTER-MODULE FLOW DISCOURAGER - An assembly for a gas turbine engine includes a first module, a second module, and a flow discourager. The second module is connected to the first module along a joint. The flow discourager is connected to the first module and extends to be received in a notch in the second module. The flow discourager acts to direct an ingestion gas flow away from the joint between the first module and the second module. | 09-11-2014 |
20140255174 | MANUFACTURE OF FULL RING STRUT VANE PACK - A vane assembly for a gas turbine engine is disclosed and includes a plurality of individual vane segments forming an annular ring. Each of the plurality of vane segments is formed of a ceramic matrix composite including at least one airfoil extending between an inner platform and an outer platform. An outer wrap is disposed about an outer periphery of the annular ring and over an interface between outer platforms of adjacent ones of the plurality of vane segments. | 09-11-2014 |
20140255202 | TURBINE ENGINE NOSECONE WITH DEFORMATION REGION - An assembly for a turbine engine including a nosecone that extends along an axis between a tip end and a second end. The nosecone includes a base region and a deformation region configured integral with the base region. The base region is arranged axially between the deformation region and the second end. The deformation region is adapted to deform relative to the base region upon impact of an object, and includes a stress concentrator. | 09-11-2014 |
20140271178 | Fan Blades for Gas Turbine Engines with Reduced Stress Concentration at Leading Edge - A light weight fan blade for a gas turbine engine is disclosed. The disclosed fan blade has an increased chord length at a transition section between the inner edge or single tooth attachment portion of the fan blade that connects to the rotor disk and the airfoil. Increasing the chord length at the transition section substantially reduces the stress concentration when the fan blade is operated under normal operating conditions. Reducing the stress concentration enables engineers to design lighter fan blades and larger fan blades for larger gas turbine engines. By increasing the size of gas turbine engines, additional thrust can be provided without adding additional engines to the aircraft. | 09-18-2014 |
20140283500 | THRUST EFFICIENT TURBOFAN ENGINE - A turbofan engine includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in the ambient air and a power turbine that converts the gas stream flow into shaft power. The turbofan engine further includes a propulsor section including a fan driven by the power turbine through a geared architecture at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. An Engine Unit Thrust Parameter defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about 0.15 at a take-off condition. | 09-25-2014 |
20140286748 | FAN CONTAINMENT CASE WITH THERMALLY CONFORMING LINER - A fan section of a gas turbine engine includes a fan containment case assembly includes an outer case of an aluminum alloy. The outer case extends circumferentially around an axial centerline. A thermally conforming liner assembly is located inwardly of the outer case. The thermally conforming liner assembly includes a circumferential liner of an aluminum alloy. A ballistic liner is located between the outer case and the thermally conforming liner assembly. | 09-25-2014 |
20140286754 | Elongated Geared Turbofan With High Bypass Ratio - A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (D | 09-25-2014 |
20140286781 | INTEGRAL FAN BLADE WEAR PAD AND PLATFORM SEAL - A fan section for a gas turbine engine includes a fan hub having a slot. A platform is supported by the fan hub. A fan blade has a root that is received in the slot. A wear pad is provided between the fan hub and the root and includes a flap integral with the wear pad to provide a seal relative to the platform. A fan blade for a gas turbine engine includes an airfoil extending from a root. A wear pad is secured to the root and has a free end providing a flap canted toward the root. | 09-25-2014 |
20140290211 | TURBINE ENGINE INCLUDING BALANCED LOW PRESSURE STAGE COUNT - A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1. | 10-02-2014 |
20140294558 | TURBOMACHINE FLUID DELIVERY MANIFOLD AND SYSTEM - An example turbomachine fluid delivery manifold provides a journal fluid path and a gear mesh fluid path that is separate from the journal fluid path. An example turbomachine lubrication system includes a geared architecture having a journal bearing and a gear. In this example, a manifold provides a portion of a first fluid path and a portion of a second fluid path separate from the first fluid path. The first fluid path is operative to communicate a first fluid from a fluid supply to the journal bearing. The second fluid path is operative to communicate a second fluid from the fluid supply to a gear mesh. | 10-02-2014 |
20140311121 | PULSE DETONATION ENGINE HAVING A SCROLL EJECTOR ATTENUATOR - The engine ( | 10-23-2014 |
20140311149 | GAS TURBINE ENGINE FAN BLADE AIRFOIL PROFILE - A fan blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by pressure and suction sides to provide an exterior airfoil surface that extends in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by the local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location. | 10-23-2014 |
20140314540 | GAS TURBINE ENGINE SYNCHRONIZATION RING - A synchronization ring for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a synchronization ring for a variable vane assembly. A plurality of rollers are attached to the synchronization ring for engaging a case on the gas turbine engine. | 10-23-2014 |
20140314541 | TURBOMACHINE THRUST BALANCING SYSTEM - An example turbomachine thrust balancing system includes a member coupled in rotation with a turbine for transferring rotational power therefrom. A load carrying device rotatably supports the member. The load carrying device is configured to counteract substantially all of the thrust load generated by the turbine. | 10-23-2014 |
20140314542 | GAS TURBINE ENGINE EXHAUST DIFFUSER WITH MOVABLE STRUTS - An exhaust diffuser for a gas turbine engine gas turbine engine includes a multiple of circumferentially spaced struts that extend between an inner gaspath wall and an outer gaspath wall, at least one of the multiple of circumferentially spaced struts movable about a strut axis. | 10-23-2014 |
20140314546 | TURBINE ENGINE GEARBOX MOUNT WITH MULTIPLE FUSE JOINTS - A turbine engine assembly includes a turbine engine case, a gearbox and a plurality of mounts connecting the gearbox to the case. Each of the mounts includes a linkage and a fuse joint. The linkage extends towards the case substantially in a first direction. The fuse joint is configured to substantially prevent movement between the gearbox and the case when the fuse joint is subject to loading less than a threshold, and to permit a constrained movement between the gearbox and the case when the loading is greater than the threshold. The fuse joint of a first of the mounts has a different configuration than the fuse joint of a second of the mounts. | 10-23-2014 |
20140318133 | THRUST EFFICIENT TURBOFAN ENGINE - A turbofan engine according to an exemplary embodiment of this disclosure, among other possible things includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in ambient air, a power turbine converting the gas stream flow into shaft power, the power turbine rotating at a first relative rotational speed, a speed reduction device driven by the power turbine, and a propulsor section including a fan driven by the power turbine through the speed reduction device at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path, wherein an Engine Unit Thrust Parameter (“EUTP”) defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first relative rotational speed of the power turbine is less than about 0.15 at a take-off condition. | 10-30-2014 |
20140338351 | GAS TURBINE ENGINE COMBUSTOR WITH TAILORED TEMPERATURE PROFILE - A method of tailoring a combustor flow for a gas turbine engine includes controlling an airflow into a swirler to be generally uniform and controlling an airflow into a quench zone to provide a desired pattern factor. | 11-20-2014 |
20140338360 | BLEED PORT RIBS FOR TURBOMACHINE CASE - An example turbomachine structure includes a case having a radially extending port, and at least one rib configured to direct flow through and past the port. | 11-20-2014 |
20140348640 | THRUST EFFICIENT TURBOFAN ENGINE - A turbofan engine according to an exemplary embodiment of this disclosure, among other possible things includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in ambient air. A power turbine converts the gas stream flow into shaft power. The power turbine rotates at a first rotational speed. A speed reduction device is driven by the power turbine. A propulsor section includes a fan driven by the power turbine through the speed reduction device at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. The fan includes a tip diameter greater than about fifty (50) inches and an Engine Unit Thrust Parameter (“EUTP”) defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about 0.30 at a take-off condition. | 11-27-2014 |