PRATT & WHITNEY CANADA CORP. Patent applications |
Patent application number | Title | Published |
20160003154 | SUPPORT LINKS WITH LOCKABLE ADJUSTMENT FEATURE - A cam-type apparatus is included in a support link between outer and inner cases of gas turbine engines for centering the cases one to another. The cam-type apparatus is lockable to allow locking an adjusted position of the cam-type apparatus. | 01-07-2016 |
20150354818 | MULTIPLE VENTILATED RAILS FOR SEALING OF COMBUSTOR HEAT SHIELDS - A seal for sealing a combustor heat shield against an interior surface of a combustor shell, the seal comprising: an upstream rail and an downstream rail defining an intermediate groove therebetween, each rail having a sealing surface with a plurality of slots extending between an upstream wall surface and a downstream wall surface, the sealing surface conforming to the interior surface of the combustor shell and defining a leakage gap therebetween. | 12-10-2015 |
20150345789 | COMBUSTOR HEAT SHIELD - A heat shield for a combustor of a gas turbine engine has a heat shield adapted to be mounted to a combustor wall with a back face of the heat shield in spaced-apart facing relationship with an inner surface of the combustor wall to define an air gap. Rails extend from the back face of the heat shield across the air gap. Grooves are defined in at least one of the rails. The rail grooves are in fluid flow communication with the air gap when the heat shield is mounted to the combustor wall. | 12-03-2015 |
20150345294 | METHOD OF BALANCING A SPOOL OF A GAS TURBINE ENGINE - A method of balancing a spool of a gas turbine engine, the spool including a forward rotor assembly and an aft rotor assembly separated by a spacer, the method comprising: balancing the forward rotor assembly and the aft rotor assembly independently from one another using conventional balancing techniques; and mathematically modeling the spool; applying a modeled axial load to the modeled spool to determine a trim weight which would balance the spool when subjected to the modeled axial load; and physically placing the trim weight to one of the forward rotor assembly and the aft rotor assembly. | 12-03-2015 |
20150321309 | METHOD OF MACHINING A SHROUD AND GRINDING WHEEL THEREFOR - A method of machining a turbine shroud segment including inserting an annular flange of a grinding wheel through a first gap defined between the shroud retention elements and into a second gap defined between the shroud platform and an axially extending, radially inwardly facing arcuate inner surface of one of the retention elements. The wheel flange is inserted with its supporting leg and the wheel body remaining out of contact with the shroud segment and with the wheel flange remaining out of contact with the platform. The inner surface is ground through contact with an annular outer grinding surface of the wheel flange. The wheel leg and body remain out of contact with the shroud segment and the wheel flange remains out of contact with the platform during grinding. | 11-12-2015 |
20150308286 | FRANGIBLE MOUNTING ARRANGEMENT AND METHOD FOR PROVIDING SAME - A frangible mounting arrangement between a bearing and a bearing support in a gas turbine engine comprises a plurality of frangible bolts connecting mounting flanges of the bearing and the bearing support. The plurality of frangible bolts is disposed on a circle. A distance on the circle between a first pair of adjacent frangible bolts is greater than a distance on the circle between a second pair of adjacent frangible bolts. The frangible bolts are resistant to axial loads and being configured to break when subjected to a breaking load above a predetermined value. The breaking load results from at least one of a bending moment and a shear load on the mounting flanges. When subjected to the breaking load, the first pair of adjacent frangible bolts breaking before the second pair of adjacent frangible bolts. A method of providing a frangible mounting arrangement between a bearing and a bearing support in a gas turbine engine is also provided. | 10-29-2015 |
20150292854 | METHOD OF DETERMINING A RADIUS OF A CUTTING END OF A TOOL FOR A TURNING MACHINE - A method of determining a radius of a cutting end of a tool for a turning machine using a touch probe is provided. One of the cutting end and the touch probe is movable relative to a reference frame having a first axis and a second axis and having a reference point trackable in the reference frame. The method comprises establishing a first contact point and recording a first coordinate of the reference point on the first axis; establishing a second contact point and recording a second coordinate of the reference point on the second axis; establishing a third contact point and recording a third coordinate of the reference point on the first axis and a fourth coordinate of the reference point on the second axis upon contact; and determining a radius of the cutting end based on the first, second, third and fourth coordinates. | 10-15-2015 |
20150291286 | MULTIPLE AIRCRAFT ENGINE CONTROL SYSTEM AND METHOD OF COMMUNICATING DATA THEREIN - The multiple aircraft engine control system having a corresponding engine controller associated with each one of the engines, each one of the engine controllers having at least two independent channels, each one of the at least two independent channels having at least two communication buses, each one of the at least two communicating buses of each channel being connected to a respective one of the at least two communicating buses of each one of the other channels. The method can time-interweave originating data of the channels. | 10-15-2015 |
20150290756 | METHOD OF MACHINING A PART - The method can include: assembling the part to be machined to a pallet into an assembly; measuring a centering of the part relative to the pallet; adjusting the centering of the part relative to the pallet based on said measuring; repeating the steps of measuring and adjusting until determining that the part is centered based on said measuring; and mounting the centered assembly to an automatic machining device and machining the part of the centered assembly with the automatic machining device. | 10-15-2015 |
20150275916 | COMPRESSOR VARIABLE VANE ASSEMBLY - A variable vane assembly for a gas turbine engine compressor and method of manufacturing same is described. A plurality of projections the inner and/or outer shroud which protruded into the annular gas path such as to ensure that a radial clearance gap, defined between the projections and a vane airfoil overhang portion, remains substantially constant throughout a substantial portion of the vane pivot arc of the variable vane. The method includes forming one cavities within the shroud, the cavities isolated from the annular gas path and disposed radially beneath at least each of the projections, and providing one or more structural reinforcing elements within the cavities. | 10-01-2015 |
20150275764 | SHAFT ASSEMBLY OF A GAS TURBINE ENGINE AND METHOD OF CONTROLLING FLOW THEREIN - A gas turbine engine comprises a shaft assembly including a hollow shaft of the gas turbine engine and a plug connected to the inlet end of the shaft. The hollow shaft has a shaft bore having a bore diameter. The hollow shaft has an inlet end for receiving a first portion of an incoming air flow. The plug has a plug bore therethrough, and an inlet end having an inlet diameter. The inlet diameter of the plug is smaller than the bore diameter. The plug includes a deflection surface adapted to deflect a second portion of the incoming air flow away from the shaft bore. A plug for connecting to an end of a hollow shaft of a gas turbine engine and s method of controlling a flow of fluid through a shaft having a bore therethrough of a gas turbine engine are also presented. | 10-01-2015 |
20150273640 | POSITIONING ASSEMBLY AND METHOD - A positioning assembly for positioning a rotary part to be machined includes a machining fixture and a positioning device having an elastically-deformable diaphragm concentrically mounted to the fixture and defining an annular loading zone which receives an axial force substantially parallel with a center axis of the fixture. The positioning device has circumferentially-spaced engaging segments, fixed to the diaphragm and extending away therefrom, which have contact members that are displaced radially to frictionally engage the rotary part when the axial force is applied to the loading zone of the diaphragm. | 10-01-2015 |
20150260103 | INTEGRATED STRUT AND IGV CONFIGURATION - A strut and IGV configuration in a gas turbine engine positioned at an upstream of a rotor includes a plurality of radial struts, for example for bearing engine loads, and a plurality of inlet guide vanes positioned axially spaced apart from the struts. The number of inlet guide vanes is greater than the number of struts. The struts are circumferentially aligned with the inlet guide vanes. | 09-17-2015 |
20150251770 | SYSTEM AND METHOD FOR OPERATING A MULTI-ENGINE AIRCRAFT IN AN AUXILIARY POWER UNIT MODE - A first and second engine are connected to a drive train for driving an aircraft accessory. A gearbox is connected to a primer mover propulsor and an actuator operatively associated with a selected engine is moveable between a position in which the selected engine drivingly engages the gearbox for driving the propulsor and a position in which the selected engine disengages from the gearbox. A position signal, a status signal, and a request signal respectively indicative of a present position of the actuator, a governing state and present speed of each engine, and a request for movement of the actuator from the present position to the other position are received. If the selected engine's speed differs from a predetermined threshold, a control signal is output for causing the engine's speed to be adjusted towards the threshold. A control signal indicating that movement of the actuator is permitted is then output. | 09-10-2015 |
20150247641 | COMBUSTION SYSTEM FOR A GAS TURBINE ENGINE AND METHOD OF OPERATING SAME - A gas turbine engine comprises a combustion system comprising a secondary annular combustor and a primary combustor in fluid communication with the secondary combustor, a secondary fuel injector associated with the secondary combustor, a primary fuel injector associated with the primary combustor, and a ECU controlling fuel delivery to the secondary and primary fuel injectors. The primary fuel injector delivers fuel to the primary combustor. The ECU allows fuel to be delivered to the secondary fuel injector in addition to the primary fuel injector only when a fuel amount higher is requested delivered by the primary fuel injector. A method of operating a gas turbine engine is also presented. | 09-03-2015 |
20150247420 | FLAPPER VALVE ASSEMBLY AND METHOD OF FLOWING AIR THERETHROUGH - A valve assembly for a gas turbine engine comprising two flappers movable independently from one another between an open position and a closed position; and a fairing disposed between the two flappers. The fairing has a portion extending downstream of the two flappers. The portion is a downstream portion of a substantially streamlined body. The two flappers are movable relative to the fairing, and when in the open position, the two flappers abut the fairing and form an upstream portion of the substantially streamlined body. A method of flowing air through a valve assembly for a gas turbine engine is also presented. | 09-03-2015 |
20150247419 | TURBINE BLADE FOR A GAS TURBINE ENGINE - A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords extending between a leading edge and a trailing edge of the airfoil. A generally round dimple is disposed on the pressure side. The dimple is contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge. The dimple is configured to initiate fracture of the blade at a predetermined speed of rotation. A method of preventing rupture of a disk of a turbine rotor is also presented. | 09-03-2015 |
20150247407 | POWER TURBINE BLADE AIRFOIL PROFILE - A power turbine includes a first stage blade having an airfoil with a cold un-coated nominal profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. | 09-03-2015 |
20150246371 | METHOD OF COLD SPRAYING COMPONENTS OF A GAS TURBINE ENGINE MASK THEREFOR - A method of cold spray coating a target surface of a component, the coating provided using selected solid powders, the method comprising: placing a mask onto the component to cover an area of the component adjacent the target surface which is not to be coated, the mask having a masking top surface provided of a material selected to be non-adhesive with the selected solid powders when cold-sprayed onto the masking top surface, the mask having a melting point above a temperature at which cold spray is performed; cold spraying the target surface with the selected solid powders, including at least some overspraying onto the mask; removing the overspray from the mask; and removing the mask from the component. A mask for a cold sprayed component of a gas turbine engine is also presented. | 09-03-2015 |
20150241309 | COMPONENT OF A GAS TURBINE ENGINE AND METHOD OF DETECTING A CRACK THEREIN - A component of a gas turbine engine comprises a substrate, a corrosion resistant top layer, and an intermediate corrodible layer disposed between the corrosion resistant top layer and the substrate. When corroding, the intermediate layer has a color contrasting with a color of the top layer. A method of detecting a crack when it penetrated the top layer in a component of a gas turbine engine having a corrosion resistant top layer and an intermediate corrodible layer comprises, in sequence, observing that at least one area of the component has a color contrasting with that of the top layer; determining that the color of the at least one area is a result of corrosion of the intermediate corrodible layer; and determining that the top layer has a crack as a result of determining corrosion of the intermediate layer. A method of facilitating crack detection in a component is also presented. | 08-27-2015 |
20150240976 | MOVEMENT-CONSTRAINING ASSEMBLY FOR FLUID-CONVEYING SYSTEM - A movement-constraining assembly for a fluid-conveying system comprises a fluid-conveying tube defining an inner passage for fluid to pass therethrough. The tube is adapted to be connected to components of the fluid-conveying system at opposed ends thereof. A blocking ring is mounted to the tube with complementary surfaces between the tube and the blocking ring to block rotation between the tube and the blocking ring, the blocking ring having a first joint portion. A base is adapted to be secured to a structure, and having a second joint portion operatively joined to the first joint portion of the blocking ring to form a joint blocking at least an axial rotational degree of freedom of the fluid-conveying tube and allowing at least one translational degree of freedom of the tube relative to the structure. A method for constraining movement of a fluid-conveying tube of a fluid conveying-system is also provided. | 08-27-2015 |
20150240721 | AIRCRAFT COMPONENTS WITH POROUS PORTION AND METHODS OF MAKING - A component including a porous portion which may be permeable or impermeable to air, and a method for making. In one example, the component is a cooled wall segment for a gas turbine engine, including a body defining a contact surface configured to be in contact with circulating hot gas and an outer surface configured to be in contact with cooling air. The body includes a first portion with at least one retention element, and a porous second portion made of a porous material permeable to air, containing a plurality of interconnected pores, and having a porosity greater than that of the first portion. The second portion is engaged to the first portion, defines at least part of the contact surface, and defines at least part of a fluid communication between the outer surface and the contact surface through the interconnected pores. The wall segment may be for example a heat shield or shroud segment. Methods of forming components are also discussed. | 08-27-2015 |
20150239010 | METHOD OF FORMING AN ABRADABLE COATING FOR A GAS TURBINE ENGINE - A method of forming an abradable coating on a gas turbine engine component comprising, in sequence: placing dry lubricant particles and trapping particles in a channel having a spraying end and containing a gas; causing at least one shockwave in the gas to travel in the channel toward the spraying end, the at least one shockwave causing the dry lubricant particles and the trapping particles to travel in the channel with it, the at least one shockwave reducing interparticle spacing and increasing particles density; directing a resulting flow of the dry lubricant particles and the trapping particles from the spraying end at a supersonic velocity to impact the component; and then plastically deforming the trapping particles upon impacting the component with the resulting flow thereby trapping the dry lubricant particles with the deformed trapping particles onto the component to provide the abradable coating. | 08-27-2015 |
20150226232 | CENTRIFUGAL COMPRESSOR DIFFUSER AND METHOD FOR CONTROLLING SAME - A centrifugal compressor having at least a diffuser is disclosed. The diffuser has an annular diffuser body having circumferentially spaced apart diffuser passages defining fluid paths through the diffuser body. The diffuser also has a plurality of diffusion members mounted to the annular diffuser body. Each diffusion member has a member inlet in fluid communication with a diffuser passage and a member outlet. Each diffusion member defines an aerodynamic throat disposed between the member inlet and the member outlet. The diffuser also has a fluid injection assembly with multiple injection conduits. Each injection conduit extends between a conduit inlet configured to receive a flow of compressible fluid from a supply and a conduit outlet communicating with a corresponding diffusion member downstream of the aerodynamic throat. The compressible fluid is injected through the conduit outlet into the diffusion members. | 08-13-2015 |
20150226070 | SHROUDED BLADE FOR A GAS TURBINE ENGINE - A turbine blade for a gas turbine engine comprises a platform, a blade root, an airfoil portion defining pressure and suction sides, and a shroud provided at a tip of the airfoil portion opposite to the blade root. The shroud includes a body having a radially outer face opposite to the airfoil portion, upstream and downstream generally parallel fins extending outwardly from the outer face, and two ribs extending outwardly from the outer face. Each of the fins has an end disposed toward the pressure side and another end disposed toward the suction side. The ribs extend from and connecting the upstream fin to the downstream fin at locations other than the ends of the upstream and downstream fins. The ribs converge toward the downstream fin at an angle of between about 10 and about 45 degrees. A shroud for a blade is also presented. | 08-13-2015 |
20150218963 | COOLING SYSTEM AND METHOD FOR SUPPLYING A COOLING GAS FLOW - A turbine case cooling system and a method for supplying a cooling gas flow are provided. The cooling system has a turbine case and a turbine case cooling manifold. The cooling system also has a fluid or cooling conduit. The cooling conduit has an inlet in fluid communication with the bypass duct, and an outlet in fluid communication with the cooling manifold. The cooling conduit also has an ejector section which in use supplies a motive air flow radially into the cooling conduit to draw a bypass air flow from the bypass duct. The motive air flow mixes with the bypass air flow to form the cooling gas flow. The cooling conduit also has a diffuser section which in use conveys the cooling gas flow toward the outlet in a direction substantially perpendicular to the center axis of the gas turbine engine. | 08-06-2015 |
20150218704 | METHOD FOR APPLYING A COATING TO A SUBSTRATE - A method for applying a coating to a substrate surface is provided. The method involves cold spraying a coating material against the surface of the substrate at a first velocity. The first velocity is lower in magnitude than a critical velocity. The method also involves cold spraying the coating material against the surface of the substrate at a second velocity. The second velocity is greater in magnitude than the critical velocity. The critical velocity is a threshold velocity below which the coating material is substantially deflected by the surface of the substrate and above which the coating material substantially adheres to the surface of the substrate. | 08-06-2015 |
20150211546 | MULTISTAGE AXIAL FLOW COMPRESSOR - A multi-stage axial compressor with an inner wall including a step portion for each of the compressor stages. Each step portion is defined along a respective stage. Each step portion may extend over at least a majority of an axial length of the stage. Each step portion may optionally include a point aligned with a maximum thickness of the airfoil portions of the rotor blades and a point aligned with a maximum thickness of the stator vanes. Adjacent step portions are connected by a transition portion converging toward a central axis of the compressor from the upstream step to the downstream step. Each transition portion has a steeper slope than that of the adjacent step portions. A method of directing flow through a multi-stage axial flow compressor is also discussed. | 07-30-2015 |
20150211545 | SHROUD TREATMENT FOR A CENTRIFUGAL COMPRESSOR - The centrifugal compressor described includes an impeller shroud which encloses the impeller and has a curved shroud surface that extends between an inducer portion and an exducer portion. The compressor includes one or more circumferential grooves in the shroud body within the exducer portion. Each groove has opposed wall segments spaced apart therefrom. The wall segments are inclined at a nonzero groove angle relative to a normal of the shroud surface in a direction opposite the fluid flow path along the shroud surface. | 07-30-2015 |
20150211540 | BLEED VALVE - A bleed valve having a piston with a sealing member, the piston being displaceable between a first position to seal an aperture of a fluid conduit and an opposed second position spaced apart from the aperture to allow the working fluid therethrough. The bleed valve also has a guiding assembly with at least one guide wheel mounted to the piston and being displaceable therewith along a guide rail. The guide wheel has a guide groove extending inwardly from an outer rolling surface. The guide rail has a rail surface with a guide protrusion. The guide groove and the guide protrusion are complementary and in rolling contact with one another. The bleed valve also has a displacement mechanism for displacing the piston between the first and second positions. A compressor with a bleed valve and method for controlling bleeding of a working fluid are also discussed. | 07-30-2015 |
20150211420 | COMBUSTOR IGNITER ASSEMBLY - A gas turbine engine comprising a combustor having a combustor liner assembly and a mounting bracket provided on the combustor liner assembly, a floating collar being slidingly received on the mounting bracket for relative sliding movement in a plane normal to an axis of an igniter opening in the liner assembly. The floating collar includes an annular surface defining a collar opening, and an igniter having an axis concentric with the axis of the collar opening is sealingly engages the annular surface. A plurality of purge openings defined in at least one of the igniter and the floating collar form cooling airflow passages communicating from the plenum to the cavity. | 07-30-2015 |
20150209054 | CUTTING TOOL AND CORRESPONDING ASSEMBLY - A one-piece cutting tool has a shank portion with a mounting end and a cutting end. The cutting end has a body portion and a substantially hemispherical portion. The cutting end further comprises a plurality of flutes, each flute helically extending adjacent to one another about the cutting end along a length from the body portion to the hemispherical portion. Each flute has a cutting edge divided into a first cutting edge section and a second cutting edge section, the first cutting edge section extending along the body portion and being substantially straight-edged or serrated, and the second cutting edge section extending along the hemispherical portion and being serrated along at least a part of the hemispherical portion. A cutting tool assembly is also provided. | 07-30-2015 |
20150204439 | VARIABLE VANE ACTUATING SYSTEM - A variable vane assembly for a gas turbine engine having an actuating system including a rotatable face gear and a respective pinion engaged to and extending transversely from the end of each of the moveable vanes. The teeth of each pinion define land surfaces angled with respect to adjacent ones of the land surfaces of the teeth of the face gear meshed therewith. A smallest axial distance between the adjacent land surfaces of the meshed pinion and face gear teeth define a backlash of the actuating system. At least one shim has a thickness adjusting an axial distance between the pinion and the face gear to set the backlash to a predetermined value. An engine with a compressor with a variable vane assembly and a method of adjusting angular variance in an actuating system for variable vanes are also discussed. | 07-23-2015 |
20150202742 | GRINDING WHEEL AND METHOD - A grinding wheel including an impeller structure for circulating coolant within the wheel defining a plurality of grooves. An angle between a second wall of each groove and a tangent to the outer circumference adjacent the groove outlet is defined based on the desired coolant flow rate, the predetermined wheel rotational speed, the radius of the outer circumference of the impeller structure, the combined surface area of the groove outlets, and a desired angle of exit of the coolant being at most 15 degrees. Also, a wheel is disclosed where the angle of the second wall of each groove is defined based on the desired coolant flow rate, the combined surface area of the groove outlets, the tangential speed of the wheel, and a value of n being at least 0.9 and less than 1. A method for distributing a coolant to a grinding site is also discussed. | 07-23-2015 |
20150198627 | SYSTEM AND METHOD FOR SPEED SENSOR POSITION DETECTION IN A MULTIPLE CHANNEL CONTROL SYSTEM - A system and method for detection of speed sensor position in an engine comprising speed sensors and a controller having channels each operatively connected to a different one of the speed sensors. A memory stores predetermined sensor position arrangements each identifying, for each one of the channels, a speed sensor connected to the channel, and predetermined engine parameter values each associated with a corresponding predetermined sensor position arrangement. A current engine parameter value is computed on the basis of received input data. The predetermined sensor position arrangements and the predetermined engine parameter values are retrieved from the memory and a predetermined engine parameter value that matches the current engine parameter value is determined. The predetermined sensor position arrangement associated with the predetermined engine parameter value is then identified and, for each one of the channels, the speed sensor currently connected to the channel is determined. | 07-16-2015 |
20150198096 | INTERNAL MANIFOLD WITH FUEL INLET - An internal fuel manifold assembly includes a fuel manifold ring having at least one fuel conveying passage in fluid flow communication with at least one inlet passage defined through an inlet member. The inlet member is connected to the manifold ring proximate its first end and connected proximate its second end to at least one transfer tube. The inlet member includes a drainage passage for collecting possible leaked fuel from an annulus, defined between the inlet member and a heat shield surrounding the inlet member and spaced apart therefrom. Any leaked fuel is discharged out of the inlet member at an exit of the fuel passage on an end surface of the inlet member proximate the second end thereof. | 07-16-2015 |
20150198091 | ELECTRIC PROBE ASSEMBLY, GAS TURBINE ENGINE HAVING SAME AND METHOD OF COOLING SAME - A gas turbine engine comprises a hot section module which includes at least one electric probe assembly surrounded by an immediate environment having a local ambient temperature. The at least one electric probe assembly includes an electric probe having a probe body, and a conduit surrounding at least a portion of the probe body. In operation, the conduit carries a fluid flowing at a temperature lower than the local ambient temperature. The at least portion of the probe body is embedded in the fluid inside the conduit. The fluid thermally insulates the probe body from the local ambient temperature of the immediate environment around the at least one electric probe assembly. | 07-16-2015 |
20150192069 | RECUPERATOR FOR GAS TURBINE ENGINE - A recuperator inserted in the exhaust duct of a gas turbine engine includes a casing surrounding a core having spiral cross channels. Inlet and outlet openings are defined in the casing for the passage of hot exhaust gases through the exhaust channels in the core. Feeder members extend radially across the outlet opening for passing the pre-combustion stage air through the air channels of the core, and header members extend radially across the inlet opening of the casing for receiving and redirecting air from the air channels in the core towards the combustor section. The feeder and header members each have a tapered configuration from the casing to the axial center of the casing so as to maintain a relatively constant pressure over the radial extent of the respective feeders and headers. | 07-09-2015 |
20150177116 | MAGNETIC CHIP DETECTOR / COLLECTOR - A magnetic chip device such as a chip detector or chip collector, comprising a horseshoe magnet held within a housing and having two adjacent tips protruding from the housing, and having a plurality of portions extending sequentially from one of the tips to the other, the plurality of portions including two end portions, each one of the two end portions extending to a corresponding one of the tips, and a rare-earth magnet portion made of a rare earth magnet material and positioned between the two end portions in the sequence. | 06-25-2015 |
20150177091 | DEVICES AND METHODS FOR BALANCING A HIGH-PRESSURE SPOOL OF A GAS TURBINE ENGINE - Devices and methods useful for balancing high-pressure spools of gas turbine engines are disclosed. An exemplary device may comprise: an input shaft configured to be coupled to an output of an accessory gear box driven by a high-pressure spool of a gas turbine engine; a first trigger rotatably coupled to the input shaft at a first speed ratio; and a sensor configured to detect the trigger at each revolution of the trigger. The first speed ratio may permit a rotational speed of the first trigger to be substantially the same as a rotational speed of the high-pressure spool. Upon detection of the trigger, the sensor may generate one or more signals representative of each associated revolution of the high-pressure spool of the gas turbine engine. | 06-25-2015 |
20150176492 | OIL TANK AND SCAVENGE PIPE ASSEMBLY OF A GAS TURBINE ENGINE AND METHOD OF DELIVERING AN OIL AND AIR MIXTURE TO SAME - An oil tank and scavenge pipe assembly of a gas turbine engine comprises a tank, and a scavenge pipe having a discharge portion disposed inside the tank. The discharge portion comprises a first portion having first and second ends. The first end is adapted to connect to an oil return line for receiving a mixture of oil and air. A bend extends from the second end downstream thereof relative to a flow of the mixture of oil and air through the scavenge pipe. The bend is configured to cause stratification of the mixture of oil and air as the mixture of oil and air flows through it. An outlet downstream of the bend delivers the mixture of oil and air to the tank. A method of delivering an oil and air mixture to a rotating oil volume of a tank of a gas turbine engine is also presented. | 06-25-2015 |
20150176486 | GAS TURBINE ENGINE WITH TRANSMISSION - A gas turbine engine with a transmission having a variable ratio is discussed. A first gear train is in driving engagement with the low pressure spool and has a first output gear. A second gear train is in driving engagement with the high pressure spool and has a second output gear spaced apart from the first output gear. A third gear train defines a driving engagement between the low pressure spool and the low pressure compressor rotor with a variable transmission ratio. A fourth gear train is in driving engagement with the first and second output gears, and in driving engagement with the third gear train to determine the transmission ratio. A method of adjusting a speed of a low pressure compressor rotor of a gas turbine engine is also discussed. | 06-25-2015 |
20150176432 | GAS TURBINE CASE AND REINFORCEMENT STRUT FOR SAME - A case assembly for a gas turbine engine comprising annular case components each having a central axis. Radial struts each have a radial axis and intersect the annular case components. A stress dissipation mass projecting from a continuous surface of at least one of the struts at the intersection with a corresponding annular case component, the stress absorption mass being on either side of a plane passing through the radial axis of the strut and the central axis of the corresponding annular case component. A method for dissipating thermal and mechanical stresses on a strut in a case assembly for a gas turbine engine is also provided. | 06-25-2015 |
20150176431 | DOUBLE FRANGIBLE BEARING SUPPORT - A double frangible bearing support structure supports a low pressure rotor of an aircraft engine. The support structure has a first bearing assembly including a first bearing supported by a first bearing support adapted to buckle or frange when subject to a predetermined critical load resulting from an abnormal rotor imbalance. The support structure has a second bearing assembly comprising a second bearing having rolling elements disposed between inner and outer races. The outer race is connected to a second bearing support by means of frangible bolts adapted to fail when subject to a predetermined critical load resulting from radial displacements and loads of the low pressure rotor following decoupling/franging at the first bearing support. | 06-25-2015 |
20150176427 | POST FBO WINDMILLING BUMPER - A bearing arrangement rotatably supports a shaft of an aircraft engine. The bearing arrangement comprises a bearing having rolling elements disposed between inner and outer races. The inner race is affixed to the shaft. A decoupler normally structurally couples the outer race of the bearing to a stator structure of the engine. The decoupler is configured to release the bearing from the stator structure when subject to a predetermined critical load. A bumper is mounted to the stator structure and encircles the bearing. The bumper has a radially inwardly facing surface disposed in close proximity to a radially outer surface of the outer race of the bearing and defines therewith a radial gap to accommodate and constrain an orbiting motion of the rotor about the central axis of the engine after decoupling at the bearing. The bumper further has an axially forwardly facing surface which is axially spaced by a predetermined axial fore gap from a first flange projecting radially outwardly from a front end portion of the outer race of the bearing. The first flange of the outer race is axially trapped between the stator structure and the bumper. After decoupling, the bearing is free to axially and radially move within the radial gap and the axial fore gap. | 06-25-2015 |
20150176425 | WALL CONSTRUCTION FOR GASPATH TRAVERSING COMPONENT - A gaspath traversing component of a gas turbine engine comprises a wall having an outer edge surface and a thickness relative to the gaspath, the wall having a plurality of layers of composite materials forming the thickness. A wear indication layer is embedded within the plurality of layers of composite material, the wear indication layer being visually contrasting with the composite material. The wear indication layer is positioned interiorly of at least one layer of said plurality of layers of composite material relative to the outer edge surface. A method for attending to a gas traversing component of a gas turbine engine is also provided. | 06-25-2015 |
20150176418 | COMPRESSOR VARIABLE VANE ASSEMBLY - A variable vane assembly for a gas turbine engine compressor with a plurality of pivoting variable vanes extending between inner and outer shrouds and having an overhang portion that protrudes from a button at opposed ends of the vane. A plurality of projections, disposed on at least one of the inner and outer shrouds, protrude into the gas path relative to a nominal gas path boundary of the shrouds. The projections are disposed adjacent the overhang portion and have an angled planar surface that is substantially parallel to a plane swept by a terminal edge of the overhang portion when the variable vane is rotated through its vane pivot arc, so that a radial clearance gap between the shroud and the overhang portion remains substantially constant through a substantial portion of the vane pivot arc. | 06-25-2015 |
20150174610 | METHOD OF SPRAY COATING A SURFACE HAVING A MAGNESIUM BASE - A method of spray coating a surface having a magnesium base is provided. The method includes, in sequence, applying a magnesium oxidizing agent onto the surface; determining whether an entirety of the surface has oxidized as a result of applying the magnesium oxidizing agent onto the surface; and spray coating the surface. | 06-25-2015 |
20150135720 | COMBUSTOR DOME HEAT SHIELD - A combustor heat shield has lips with fins distributed on the lips. The lip-fins have an extended end portion projecting rearwardly from the back face of the heat shield. Impingement jets may be directed against the rearwardly extended end portions of the lip-fins to enhance cooling. The heat shield may define a fuel nozzle opening surrounded by a rail on the back side of the heat shield. Impingement holes or slots may be defined in the rail for allowing cooling air passing therethrough to impinge upon the lip-fins. | 05-21-2015 |
20150118023 | ACOUSTIC STRUCTURE FOR A GAS TURBINE ENGINE - An acoustic structure for a gas turbine engine comprising an noise reduction layer and a fire protector layer connected to the noise reduction layer. The noise reduction includes a perforated inner wall adapted to be in contact with a first fluidic environment, and an noise reduction adjacent to the inner wall. The fire protector layer includes a non-perforated outer wall adapted to be in contact with a second fluidic environment having potentially a fire, a fire protector adjacent to the outer wall, and a pressure resisting wall disposed between the fire protector and the noise reduction. The second fluidic environment is under a pressure lower than a pressure of the first fluidic environment. The inner and outer walls are load-bearing walls of the acoustic structure | 04-30-2015 |
20150113994 | COMBUSTOR FOR GAS TURBINE ENGINE - In a gas turbine combustor having an inner and outer liner defining an annular combustion chamber, at least an annular scoop ring provided on each inner and outer combustor liner. The annular scoop ring includes a solid radial inner base provided with bores defined therein and communicating with the combustion chamber to form air dilution inlets. The scoop ring has a radial outer portion in the form of a C-shaped scoop open to receive high velocity annular air flow. The bores of the inlets communicating with the scoop portion to direct the air flow into the combustion chamber whereby the bores of the inlets form jet nozzles to generate air jet penetration and direction within the combustion chamber. | 04-30-2015 |
20150110628 | FASTENING SYSTEM FOR ROTOR HUBS - A rotor disk assembly comprises a first rotor disk with a plurality of circumferentially distributed first throughbores. A second rotor disk comprises a connection portion projecting at least partially axially, a plurality of circumferentially distributed second throughbores being provided in the connection portion in cooperative distribution relative to the first rotor disk for the first and second throughbores to be in register with one another in the rotor disk assembly. Connector bolts each have an elongated body with a flange between its ends, a head at its first end and a removable head at its second end, the head at the first end spaced apart from the flange for the connector bolt to be secured to one of the rotor disks at a respective throughbore, the removable head at the second end being spaced apart from the flange for the second end to project. An anti-rotation feature is between each said connector bolt and at least one of the rotor disks to prevent rotation of the connector bolts when the removable head is installed on the second end in the rotor disk assembly. | 04-23-2015 |
20150107256 | COMBUSTOR FOR GAS TURBINE ENGINE - An assembly of a combustor and fuel manifold comprises a fuel manifold having at least an annular portion with a plurality of fuel outlets facing at least partially radially inward. A combustor comprises an annular inner liner and an annular outer liner spaced apart from the inner liner, the inner liner and outer liner concurrently forming an annular combustor chamber therebetween. The inner and outer liners concurrently define an annular receptacle in the annular combustor chamber for receiving the fuel manifold. The inner liner and outer liners are shaped to define an annular gas path in the annular combustor chamber having an upstream portion that is substantially radial and oriented toward a center when the combustor is in a gas turbine engine. | 04-23-2015 |
20150099453 | METHOD OF MANUFACTURING RECUPERATOR AIR CELLS - A method of manufacturing a recuperator segment uses metal tubes deformed into air cells in a waved configuration. The air cells are stacked one to another to form a double skinned recuperator segment providing cold air passages through the respective air cells and hot gas passages through spaces between adjacent air cells. | 04-09-2015 |
20150098812 | INTEGRATED STRUT AND TURBINE VANE NOZZLE ARRANGEMENT - An integrated strut and turbine vane nozzle (ISV) arrangement according to an embodiment, includes a single-piece interturbine duct (ITD) and a plurality of vane nozzle segments removably attached to the ITD. Vane airfoils of the vane nozzle segments in combination with trailing edge portions of the struts, form a vane nozzle integrated with the ITD. | 04-09-2015 |
20150096641 | FLOW REGULATING APPARATUS - A flow regulating device in a fluid system according to one embodiment may include a housing defining a cavity receiving an orifice body. The orifice body defines a bore extending axially therethrough and includes a flexible end section, with axial slits in one example, to allow the end section to be radially compressed to change a cross-section of a flow passageway when the bore of the orifice body is included in the flow passageway. | 04-09-2015 |
20150096302 | COMBUSTOR HEAT-SHIELD COOLING VIA INTEGRATED CHANNEL - A combustor heat shield for a gas turbine engine has a heat shield panel adapted to be mounted to an inner surface of a combustor shell with a back face of the panel spaced-apart from the combustor shell to define an air gap therewith. Studs project from the back face of the panel for engagement in corresponding mounting holes defined in the combustor shell. Each stud has a threaded distal end portion for engagement with a nut outside of the combustor shell. At least one of the studs has a channel defined in a peripheral surface thereof. The channel extends longitudinally along the stud from an inlet end connectable to a source of cooling air outside of the combustor shell to an outlet end disposed within the air gap for locally providing cooling air at the base of the stud. | 04-09-2015 |
20150093281 | Method of Creating a Surface Texture - A method of creating a texture on at least one surface of a part is disclosed. The part is molded from a feedstock including a powder remaining solid during molding and a binder, and solidified. Then, a physical state of the binder is changed in only a predetermined portion of each surface of the part to be textured. The texture is then created from the predetermined portion by debinding and sintering the part. | 04-02-2015 |
20150093252 | INTERNALLY COOLED AIRFOIL - An internally cooled airfoil for a gas turbine engine has a hollow airfoil body defining a core cavity. An insert is mounted in the core cavity. A cooling gap is provided between the insert and the hollow airfoil body. A plurality of standoffs project across the cooling gap. Trip-strips projecting laterally between adjacent standoffs. The trip-strips and the standoffs may be integrated into a unitary heat transfer feature. | 04-02-2015 |
20150093211 | BOLT FOR GAS TURBINE ENGINE ROTOR - The bolt has, in sequence along a bolt axis: a threaded portion, a thread run-out portion, a shank, and a head, the threaded portion having a thread with a given root radius and a given depth, the thread run-out portion connected to the shank via a thread run-out fillet having a thread run-out fillet radius, the thread run-out fillet radius being between two and six times the thread root radius. | 04-02-2015 |
20150088299 | MACHINE TOOL CERTIFICATION FOR PART SPECIFIC WORKING VOLUME - A method for machining a selected part with a machine tool, comprises obtaining a master part replicating at least a portion of a geometry of a selected part. The master part is loaded in a machine tool. A signature of the machine tool is defined by measuring at least dimensional data of the master part relative to the machine tool, the dimensional data being limited to a selected-part-specific working volume substantially smaller than a complete working volume of the machine tool. The machine tool is certified as being within tolerances to machine the selected part within the working volume, using the dimensional data of the signature. The selected part is machined from a workpiece with the machine tool. | 03-26-2015 |
20150086381 | INTERNALLY COOLED AIRFOIL - An internally cooled airfoil for a gas turbine engine has a hollow airfoil body including pressure and suction sidewalls defining a cooling passage therebetween. A combination of pedestal and trip-strips are used in the cooling passage to enhance heat transfer while minimizing the coolant pressure drop across these features. | 03-26-2015 |
20150086352 | Gas Turbine Engine Inlet Assembly and Method of Making Same - A method of fabricating an inlet assembly for a gas turbine engine, the method including defining an intake duct of the inlet assembly between first and second space apart inlet case portions, locating at least one strut across the intake duct, each strut having a proximal end made integral to the first inlet case portion and an opposed distal end engaged in a respective strut-receiving aperture defined through the second inlet case portion, while maintaining the distal end of each strut in the respective strut-receiving aperture, adjusting the relative position of the first inlet case portion and the second inlet case portion until a predetermined throat dimension of the intake duct is obtained, and locking the adjusted relative position by attaching the second inlet case portion to each strut. An inlet assembly and gas turbine engine with inlet assembly as also disclosed. | 03-26-2015 |
20150075169 | INTEGRATED TURBINE EXHAUST STRUTS AND MIXER OF TURBOFAN ENGINE - A turbine exhaust case (TEC) of a turbofan aeroengine includes a mixer for mixing exhaust gases with a bypass air stream, the TEC comprising an annular hub and an annular shroud with the mixer attached to a downstream end of the shroud, the shroud and the mixer surrounding the hub to form an annular exhaust gas duct positioned radially therebetween, a plurality of deswirling struts circumferentially spaced apart with respect to a central axis of the TEC and located within an axial length of the mixer between an upstream end and a downstream end of the mixer, the deswirling struts each having a cambered profile and extending radially across the annular exhaust gas duct and interconnecting the mixer and the hub. | 03-19-2015 |
20150059349 | COMBUSTOR CHAMBER COOLING - A gas turbine engine combustor which is cooled by a hybrid cooling apparatus and method. Heat shields cooled by impingement cooling air are present at an upstream section of the combustor liners and exhausted impingement cooling air is substantially discharged outside of the primary zone of the combustor chamber. The single-skinned downstream section of the combustor liners is cooled by effusion cooling. | 03-05-2015 |
20150052901 | INTERLOCKING COMBUSTOR HEAT SHIELD PANELS - A combustor heat shield assembly comprises a circumferential array of heat shield panels individually mounted to an inner surface of a combustor shell. Each heat shield panel has a sealing rail extending from a back side thereof and a plurality of bolted connections securely holding the heat shield panel on the combustor shell with the sealing rail in sealing contact with the inner surface of the combustor shell. Each pair of adjacent heat shield panels comprises first and second panels having adjoining lateral edges, the first panel having a boltless area on its back side at a location adjacent to its adjoining lateral edge. The second panel has a first one of its bolted connections provided adjacent to its adjoining lateral edge and in facing relationship with the boltless area of the first panel. A tab projects from the lateral edge of the second panel in overlapping relationship with at least a portion of the boltless area of the first panel for transferring a force from the first bolted connection of the second panel to the boltless area of the first panel. | 02-26-2015 |
20150052900 | ASYMMETRIC COMBUSTOR HEAT SHIELD PANELS - A combustor heat shield assembly comprises a circumferential array of heat shield panels individually mounted to an inner surface of a combustor shell. Each heat shield panel has opposed front and back faces, the back face facing the inner surface of the combustor shell and being spaced therefrom to define an air gap. The front and back faces have a perimeter including opposed lateral edges extending between opposed circumferentially extending edges. The lateral edges of adjacent heat shield panels have complementary non-linear profiles defining an asymmetric heat shield panel interface. | 02-26-2015 |
20150044032 | INTEGRATED STRUT AND VANE ARRANGEMENTS - In an integrated strut and turbine vane nozzle (ISV) configuration, lug/slot or tag/groove arrangements may be provided between an interturbine duct (ITD) of the ISV and a vane ring of the ISV such that struts of the ITD and associated vanes are angularly positioned to form integrated strut-vane airfoils, reducing mismatch at the integration. | 02-12-2015 |
20150040855 | Rotary Internal Combustion Engine with Static Oil Seal - A stator for a rotary internal combustion engine, with a body having an internal cavity. Each end wall has a scavenging cavity defined therein in fluid communication with the internal cavity through a respective scavenging opening extending through the inner surface thereof, and at least one annular oil seal groove defined in the inner surface thereof concentric with the central bore and located radially outwardly of the scavenging opening. At least one annular oil seal is received in each groove and protrudes from the end wall into the internal cavity for sealing engagement with a surface of a rotor of the engine, each seal being biased axially away from the end wall. A rotary internal combustion engine and a method of limiting radially outwardly directed oil leaks in a rotary engine are also disclosed. | 02-12-2015 |
20150040568 | COMBUSTOR FLOATING COLLAR ASSEMBLY - A gas turbine combustor floating collar for mounting an igniter or fuel nozzle to a combustor is provided with an outer periphery in a non-circular shape, for example having at least one section thereof formed with a flat surface, such as a square or triangular shape. The outer periphery of the floating collar is complementary to and completely surrounded by an inner periphery surface of a recess of a boss affixed on the combustor, thereby preventing substantial rotation of the floating collar with respect to the boss. | 02-12-2015 |
20150040377 | Method of Supporting a Part - A method of supporting a part with particulate shape retaining media, the method including placing the part on a bed of the particulate shape retaining media, fluidizing the particulate shape retaining media until the part penetrates therein, and vibrating the bed of particulate shape retaining media to compact the particulate shape retaining media around the part. The part may be a green part to be debound in a powder injection molding process. Fluidization may be performed through vibrations at a different frequency than the compaction. | 02-12-2015 |
20150036781 | HEALTH MONITORING OF IGNITERS - Methods and associated devices useful for the health monitoring of igniters of gas turbine engines are disclosed. Exemplary embodiments disclosed include the use of an accumulated spark count to an igniter to obtain an indication of wear on the igniter and/or the remaining life of the igniter. An exemplary method disclosed includes: receiving one or more signals indicative of a commanded spark count to the igniter for one or more ignition events; processing the one or more signals indicative of the commanded spark count to the igniter, determining a total accumulated commanded spark count to the igniter; and generating one or more signals indicative of at least one of an estimated wear on the igniter or an estimated remaining life of the igniter based on the total accumulated commanded spark count to the igniter. | 02-05-2015 |
20150033836 | METHODS AND APPARATUS FOR INSPECTING COOLING HOLES - Methods and apparatus for inspecting cooling holes in a wall of a combustor of a gas turbine engine are disclosed. An exemplary method disclosed may comprise: heating the wall of the combustor; directing a flow of cooling fluid through the one or more cooling holes in the wall of the combustor while the wall is being heated; acquiring a first measurement indicative of a flow rate of the cooling fluid through the one or more cooling holes; and acquiring a second measurement indicative of a cooling effectiveness provided by the cooling fluid flowing through the one or more cooling holes at the flow rate. | 02-05-2015 |
20140377075 | METHOD FOR REPAIRING A BLADE - A method for repairing a blade in a gas turbine engine comprises the steps of: isolating the damage on the airfoil of the blade; forming a cut back in the shape of elongated āDā shaped recess with a pair of fillets, a depth and a longitudinal axis of the āDā shaped recess having a length along the leading or trailing edge of the airfoil; and the fillets having a respective radius. | 12-25-2014 |
20140369814 | DIFFUSER PIPE FOR A GAS TURBINE ENGINE AND METHOD FOR MANUFACTURING SAME - A diffuser pipe for a gas turbine engine comprises a hollow pipe body including a first end, a second end fluidly connected to the first end, and at least one flattened area proximate to the second end. A ring is connected to the second end. The ring is an outlet of the diffuser pipe. At least one stiffener is disposed on the at least one flattened area. The ring and the at least one stiffener reduce vibratory stresses at the second end of the pipe body. A method of manufacturing a diffuser pipe of a gas turbine engine is also presented. | 12-18-2014 |
20140361119 | MOUNTING SYSTEM FOR MOUNTING ENGINE NACELLE COMPONENTS AND ASSOCIATED METHOD - The mounting system can be used in mounting a turbofan gas turbine engine having a nacelle including two halves, each half having a hinged end hingedly connected to the pylon and a free end lockingly engageable with the engine structure. The mounting system can include a primary connection connecting the engine structure to the pylon, and a flexible connection provided between the hinged connections and the pylon, the flexible connection being elastically deformed when a significant force is exerted in a transversal plane upon one of the halves in the closed position. | 12-11-2014 |
20140356159 | LOW HUB-TO-TIP RATIO FAN FOR A TURBOFAN GAS TURBINE ENGINE - A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R | 12-04-2014 |
20140356158 | GAS TURBINE ENGINE VANE ASSEMBLY AND METHOD OF MOUNTING SAME - The gas turbine engine vane assembly has a vane having an elongated airfoil body extending to a tip and a grommet disposed around the tip. An insert having a closed loop shape with an inner surface matingly shaped to receive the outer surface of the grommet, and an outer surface matingly shaped to be snugly received in the slot. A method of mounting such a vane assembly is also disclosed. | 12-04-2014 |
20140338346 | COMBUSTOR SKIN ASSEMBLY FOR GAS TURBINE ENGINE - A combustor assembly includes a hot skin of a combustion chamber wall having an inner face exposed to the combustion chamber and an opposite outer face, a receiving skin having a securing portion affixed to the hot skin outer face in an air-tight manner and a receiving flange, extending from the securing portion, that is offset from the hot skin outer face to form a female recess, a cold skin having a cold wall portion spaced from the hot skin and forming a cooling cavity therebetween, a securing portion extending from a first end of the cold wall portion affixed to the hot skin outer face in an air-tight manner and a male flange extending from a second end of the cold wall portion opposite the first end, the male flange snugly received in the female recess and forming a sliding engagement therebetween. | 11-20-2014 |
20140327489 | DYNAMICALLY DETECTING RESONATING FREQUENCIES OF RESONATING STRUCTURES - There is described herein a real-time scheme, implementable in software, hardware, or a combination thereof, to detect a resonating frequency of a structure from a sensed signal and dynamically set the center frequency of an adaptive compensator for effective attenuation of the resonating frequency. | 11-06-2014 |
20140318145 | HYBRID SLINGER COMBUSTION SYSTEM - There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand. | 10-30-2014 |
20140311152 | REVERSE FLOW CERAMIC MATRIX COMPOSITE COMBUSTOR - A gas turbine engine has an annular reverse-flow combustor with a combustor inner liner enclosing a combustion chamber. The inner liner having a dome portion at an upstream end of the combustor and a downstream combustor exit defined between a small exit duct portion and a large exit duct portion. At least one of the dome portion, the small exit duct portion and the large exit duct portion is made of a separately formed hemi-toroidal shell composed of a ceramic matrix composite. | 10-23-2014 |
20140304989 | ROTOR BLADE ASSEMBLY TOOL FOR GAS TURBINE ENGINE - A rotor blade assembly tool for coupling a plurality of circumferentially spaced rotor blades to a rotor disc of a turbine rotor assembly, includes a base ring with an array of circumferentially spaced resilient fingers axially extending from the base ring. The resilient fingers are configured each to radially abut a blade seal, damper or other engine component against radially inner facing surfaces of platforms of the respective blades when the blades are being seated onto the disc during an assembly procedure | 10-16-2014 |
20140297155 | AIRCRAFT POWER OUTTAKE MANAGEMENT - A system and method for controlling the operation of a gas turbine engine supplying power to an aircraft. The engine is controlled according to a reading of an amount of power drawn from the supplied power. The reading is fed directly to a control system, which issues commands for controlling engine parameters comprising an acceleration reference signal, load shedding, variable geometry positioning, and fuel flow. The control system may further issue commands for controlling the amount of power drawn. The control system may further use the reading to monitor the engine's condition. | 10-02-2014 |
20140291987 | ENGINE ARCHITECTURE USING ELECTRIC MACHINE - There is described a method for controlling an engine and a system architecture for an engine. The system architecture comprises a first electric machine having a single rotor dual stator configuration for operating as a starter-generator for the engine; a second electric machine having a single rotor dual stator configuration for operating as a motor; a dual channel motor drive unit coupled to the second electric machine; a dual channel power control unit coupled to the first electric machine and the motor drive unit; a dual channel full authority digital engine control (FADEC) coupled to the dual channel power control unit and the dual channel motor drive unit; and at least two accessories coupled to the second electric machine and driven by motive power from the single rotor of the second electric machine. | 10-02-2014 |
20140290265 | GAS TURBINE ENGINE WITH TRANSMISSION - A multi spool gas turbine engine with a differential having a selectively rotatable member which rotational speed determines a variable ratio between rotational speeds of driven and driving members of the differential. The driven member is engaged to the first spool and a rotatable shaft independent of the other spools (e.g. connected to a compressor rotor) is engaged to the driving member. First and second power transfer devices are engaged to the first spool and the selectively rotatable member, respectively. A circuit interconnects the power transfer devices and allows a power transfer therebetween, and a control unit controls the power being transferred between the power transfer devices. Power can thus be transferred between the first spool and the selectively rotatable member to change the speed ratio between the first spool and the rotatable shaft. | 10-02-2014 |
20140286746 | COMPRESSOR SHROUD REVERSE BLEED HOLES - A gas turbine engine compressor includes a rotor defining a central axis of rotation and a plurality of blades which project into an annular compressor gas flow passage, and a shroud circumferentially surrounding the rotor and having a radially inner surface adjacent to the blade tips. Bleed holes extend through the shroud adjacent the blade tips, each of the bleed holes having an inlet end disposed in the shroud radially inner surface and an outlet end disposed in a shroud radially outer surface. Bleed air removed from the annular gas flow passage flows through the bleed holes from the inlet to the outlet ends. The outlet end of each bleed hole is located circumferentially upstream of the inlet end relative to a direction of rotational flow in the annular gas flow passage driven by a direction of rotation of the rotor. | 09-25-2014 |
20140278014 | SYSTEM AND METHOD FOR ENGINE TRANSIENT POWER RESPONSE - There is provided a system and method for controlling an engine. A request signal indicative of a demand for the engine to output a required power level is first receive. A position control signal is then generated in response to the request signal. The position request signal is indicative of a first request for adjusting a present position of a variable geometry mechanism of the engine towards a commanded position to achieve the required power level. An acceleration rate control signal is further generated on the basis of the position control signal. The acceleration rate control signal is indicative of a second request for adjusting an acceleration rate of the engine in accordance with the commanded position of the variable geometry mechanism. The position control signal and the acceleration rate control signal are then output to the engine. | 09-18-2014 |
20140271158 | COMPRESSOR STATOR - Compressor stators ( | 09-18-2014 |
20140271108 | COMPRESSOR BLEED SELF-RECIRCULATING SYSTEM - A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annular bleed cavity surrounds the shroud and is in communication with outlet openings of the bleed holes. The bleed holes provide communication between said main gas flow passage and the bleed cavity. The bleed cavity includes exit passages having outlets disposed in said shroud at a second axial location which is upstream of both the first axial location and the leading edge of the blades of the rotor. Bleed air is passively bled from the main gas flow passage via the bleed holes, recirculated through the bleed cavity and re-injected back into the main gas flow passage at the second axial location. | 09-18-2014 |
20140271105 | TURBINE SHROUD SEGMENT SEALING - A segmented shroud ring surrounds a circumferential array of blades of a gas turbine engine rotor. The shroud ring has a plurality of shroud segments disposed circumferentially one adjacent to another. The circumferentially adjacent shroud segments have confronting sides defining an inter-segment gap therebetween. The inter-segment gaps are sealed by a sealing band mounted to the radially outer surface of the segmented shroud ring so as to extend across the inter-segment gaps around the full circumference of the shroud ring. Impingement jet holes may be defined in the sealing band for cooling the shroud segments. | 09-18-2014 |
20140265142 | CARBON SEAL ASSEMBLY - A carbon seal assembly comprises an annular seal runner adapted to be sealingly mounted to a shaft to rotate therewith. An annular member is secured to a structure, the annular member having an annular body and a projection extending from the body toward the seal runner in an axial direction relative to an axis of the runner. A axial gap is defined between the member and the seal runner when secured to the shaft and structure respectively, with the projecting extending into only a portion of the gap, such that the gap defines a first width portion and a second width portion, with the member and the seal runner being made of complementary materials for magnetic attraction therebetween. An annular carbon element is mounted to the seal runner to rotate therewith and positioned in the first width portion of the gap, the carbon element having an annular wear surface abutting against the annular member, a plane of the annular wear surface being axially offset from the second width portion of the gap. | 09-18-2014 |
20140261294 | INTERNAL COMBUSTION ENGINE WITH COMMON RAIL PILOT AND MAIN INJECTION - An internal combustion engine including a pilot subchamber, a pilot fuel injector having a tip in communication with the pilot subchamber, an ignition element positioned to ignite fuel within the pilot subchamber, and a main fuel injector spaced apart from the pilot fuel injector. The engine includes a common rail in fluid communication with the main fuel injector and with the pilot fuel injector and a pressure regulating mechanism in fluid communication with the common rail for regulating a fuel pressure therein. A method of combusting fuel in an internal combustion engine is also provided. | 09-18-2014 |
20140261293 | INTERNAL COMBUSTION ENGINE WITH PILOT AND MAIN INJECTION - An internal combustion engine with at least two rotatable bodies each defining at least one combustion chamber of variable volume and, for each rotatable body: a pilot subchamber, a pilot fuel injector having a tip in communication with the pilot subchamber, an ignition element positioned to ignite fuel within the pilot subchamber, and a main fuel injector spaced apart from the pilot fuel injector. The engine includes a common first fuel conduit in fluid communication with each main fuel injector, and a common second fuel conduit in fluid communication with each pilot fuel injector. First and second pressure regulating mechanisms which are settable at different pressure values from one another respectively regulate a fuel pressure in the first and second conduits. A method of combusting fuel in an internal combustion engine is also provided. | 09-18-2014 |
20140261292 | INTERNAL COMBUSTION ENGINE WITH PORT COMMUNICATION - An internal combustion engine with rotatable bodies each received in a respective internal cavity. The engine includes at least one inlet port for each internal cavity in fluid communication with the combustion chamber(s) thereof at least during their intake phase and a beginning of their compression phase. The bodies are angularly offset with the beginning of the compression phase of the combustion chamber(s) defined by each body being simultaneous with at least a beginning of the intake phase of the combustion chamber(s) defined by a different one of the bodies. A respective conduit provides a fluid communication between an inlet port for each body and an inlet port for the different one of the bodies, with each conduit being in fluid communication with a plenum for receiving pressurized air. A method of feeding air to an internal combustion engine is also provided. | 09-18-2014 |
20140260324 | TURBO-MACHINERY ROTORS WITH ROUNDED TIP EDGE - A rotor for a gas turbine engine includes a plurality of radially extending blades, each having a remote blade tip defining an outer tip surface, and a leading edge defined between opposed pressure and suction side airfoil surfaces. A shroud circumferentially surrounds the rotor, and a radial distance between an inner surface of the shroud and the outer tip surface of the blades defines a radial tip clearance gap therebetween. The tip of each of the blades has a pressure side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface, and a suction side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface. The suction side edge has a larger radius of curvature than the pressure side edge, thereby reducing the amount of tip leakage flow through the radial tip clearance gap. | 09-18-2014 |
20140260306 | ENGINE STARTING SYSTEM USING STORED ENERGY - There is described a method for and system for starting at least one engine from a twin engine installation. The starting system comprises a first engine arrangement comprising a first electric machine having a single rotor dual stator configuration, a first dual channel power control unit coupled to the first electric machine, and a first dual channel full authority digital engine control (FADEC) coupled to the first dual channel power control unit; a second engine arrangement comprising a second electric machine having a single rotor dual stator configuration, a second dual channel power control unit coupled to the second electric machine, and a second dual channel full authority digital engine control (FADEC) coupled to the second dual channel power control unit; an energy storage unit coupled to the first engine arrangement and the second engine arrangement and having at least a first super-capacitor and a second super-capacitor; and a DC to DC converter configured to receive a first voltage level from a power source, increase the first voltage level to a second voltage level, and charge the first super-capacitor and the second super-capacitor to the second voltage level. | 09-18-2014 |
20140260298 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor comprises an annular combustor chamber formed between the inner and outer liners. Fuel nozzles each have an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. A plurality of nozzle air holes are defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles. The nozzle air holes are configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber. A central axis of the nozzle air holes has a tangential component relative to the central axis of the annular combustor chamber. | 09-18-2014 |
20140260297 | COMBUSTOR FOR GAS TURBINE ENGINE - A combustor comprises an annular combustor chamber formed between the inner and outer liners, the annular combustor chamber having a central axis. Fuel nozzles are in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber. The fuel nozzles are oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. Nozzle air inlets are in fluid communication with the annular combustor chamber to inject nozzle air generally radially in the annular combustor chamber. A plurality of dilution air holes are defined through the inner and outer liner downstream of the nozzle air inlets, the dilution holes configured for high pressure air to be injected from an exterior of the liners through the dilution air holes generally radially into the combustor chamber, a central axis of the dilution air holes having a tangential component relative to the central axis of the annular combustor chamber. | 09-18-2014 |
20140260296 | SLINGER COMBUSTOR - A slinger combustor has an annular combustor shell defining a combustion chamber having a radially inner fuel inlet for receiving a spray of fuel centrifuged by a fuel slinger. The combustion chamber has a fuel atomization zone extending radially outwardly from the fuel inlet and merging into a radially outwardly flaring expansion zone leading to a combustion zone. A plurality of nozzle air inlets are defined in the fuel atomization zone of the combustor shell. The nozzle air inlets have a nozzle axis intersecting the stream of fuel and a tangential component in a direction of rotation of the fuel slinger. A plurality of dilution holes are defined in the combustor shell and have a dilution axis intersecting the combustion zone. The dilution axis of at least some of the dilution holes has a tangential component opposite to the direction of rotation of the fuel slinger. | 09-18-2014 |
20140260295 | GAS TURBINE ENGINE WITH TRANSMISSION AND METHOD OF ADJUSTING ROTATIONAL SPEED - A method of adjusting a rotational speed of the low pressure compressor rotor(s) of a gas turbine engine, including rotating the high pressure compressor rotor(s) with the high pressure turbine rotor(s) through the high pressure spool, rotating the low pressure turbine rotor(s) with a flow of exhaust gases from the high pressure turbine, rotating the low pressure spool with the low pressure turbine rotor(s), rotating a load of the engine with the low pressure spool, driving a rotation of the low pressure compressor rotor(s) with the low pressure spool through a variable transmission defining a variable transmission ratio between rotational speeds of the compressor rotor(s) and the low pressure spool, and adjusting the transmission ratio to obtain a desired rotational speed for the low pressure compressor rotor(s). A method of adjusting rotational speeds of a gas turbine engine and a gas turbine engine are also described. | 09-18-2014 |
20140260283 | GAS TURBINE ENGINE EXHAUST MIXER WITH AERODYNAMIC STRUTS - An exhaust mixer for a gas turbine engine has a plurality of circumferentially distributed alternating inner and outer lobes, and a plurality of aerodynamic struts positioned and oriented between adjacent inner lobes in the core flow passage. The struts are configured for diverting a portion of the core flow radially outwardly in troughs formed by the outer lobes on the core flow side of the mixer. | 09-18-2014 |
20140260266 | COMBUSTOR FOR GAS TURBINE ENGINE - A gas turbine engine comprises an annular combustor chamber formed between an inner liner and an outer liner. An annular upstream zone is adapted to receive fuel and air from an annular nozzle. An annular mixing zone is located downstream of the upstream zone. The mixing zone has a reduced radial height relative a downstream combustion zone of the combustion chamber, the mixing zone defined by straight wall sections. | 09-18-2014 |
20140260260 | COMBUSTOR FOR GAS TURBINE ENGINE - A gas turbine engine comprises a combustor. The combustor comprises an annular combustor chamber formed between an inner liner and an outer liner spaced apart from the inner liner. An annular fuel manifold has fuel nozzles distributed circumferentially on the fuel manifold, the fuel manifold and fuel nozzles positioned entirely inside the combustion chamber. | 09-18-2014 |
20140257597 | INTERFACE SYSTEM FOR MULTIPLE PROTOCOLS - An interface system for a plurality of different protocols. The system includes a single receiver, a processor and a single transmitter. The receiver may convert an input signal to a 1/0 signal according to a determined one of the plurality of different protocols. The processor may receive the converted logic level signal and parse the converted logic level signal into data according to the determined one protocol. The transmitter may receive the translated signal and convert the translated signal to an output signal having a format readable by an external device. | 09-11-2014 |
20140257542 | SYSTEM AND METHOD FOR POSITIONING ERROR COMPENSATION DURING MANUFACTURING OF COMPLEX-SHAPED GAS TURBINE ENGINE PARTS - A system and method for error compensation in positioning a complex-shaped gas turbine engine part during manufacturing thereof with a machine. Theoretical measurements for a plurality of control points on the part are first retrieved. Actual measurements for the control points are then acquired in a coordinate system of the machine. If an error between the actual and theoretical measurements is beyond a tolerance, a transformation matrix is computed. The transformation matrix represents a transformation to be applied to the coordinate system to adjust a pose thereof for compensating the error. The transformation matrix may be computed and applied to the coordinate system iteratively until the actual measurements are brought within tolerance. A machining program may then be generated for manufacturing the part accordingly. | 09-11-2014 |
20140256494 | LUBRICATION OIL SYSTEM FOR A REDUCTION GEARBOX - A reduction gear box is part of a gas turbine engine and includes a casing and reduction epicyclic gear stages within the casing. The reduction gear stages comprising at least an epicyclic array of gears meshing together with at least one planetary gear mounted for rotation on a gear carrier and a bearing associated therewith. A lubricating oil delivery system is provided within the casing, surrounding a portion of the gear carrier. The oil delivery system includes a conduit; a closed oil reservoir; a first metered opening communicating the conduit with the reservoir and a plurality of metered outlet openings communicating the reservoir with the gear carrier and the bearing such that the reservoir is filled with oil in normal flight operating conditions and the oil trapped in the reservoir is released to the carrier and bearing when a temporary drop in the oil system pressure occurs. | 09-11-2014 |
20140255613 | LOW ENERGY PLASMA COATING - A method of coating an aluminum alloy or magnesium alloy component, including cleaning and drying surfaces of the component to be coated; suspending a powdered coating material in a carrier gas and feeding the suspended powdered coating material through a plasma torch in a flowing gas; heating the coating material in the plasma torch to a molten or semi-molten state using a nominal power below 25 kW; and depositing the coating material with the plasma torch directly on the surfaces to be coated. The component may be made of a magnesium alloy containing at one or more of zinc, cerium and zirconium, or of an aluminum alloy containing one or more of magnesium, silicon, copper and chromium. The powder material may be made in majority of aluminum. | 09-11-2014 |
20140255197 | ROTOR BLADES FOR GAS TURBINE ENGINES - A rotor dual-blade for a gas turbine engine that has a first blade component extending radially between a root and a tip and a second blade component, separate from the first component, extending radially between a root and a tip, wherein the second blade component is downstream, in series, of the first blade component and at least the first blade component is made of metal while the second blade component is a light weight composite material. | 09-11-2014 |
20140255179 | LOW PROFILE VANE RETENTION - A gas turbine having an annular casing with a series of circumferentially spaced openings defined therethrough; a plurality of vanes extending radially inwardly though respective casing openings, an outer end of the vanes projecting radially outwardly from the casing through the respective openings and located within the respective openings by grommets, and an inner end of the vanes being mounted to an inner portion of the casing; a flexible, segmented strap extending around the annular casing, surrounding the projecting outer ends of the vanes; and a spring radially loading the flexible strap configured to apply a tension force to the flexible strap. | 09-11-2014 |
20140255159 | INTEGRATED STRUT-VANE - An integrated strut and turbine vane nozzle (ISV) has inner and outer annular duct walls defining an annular flow passage therebetween. Circumferentially spaced-apart struts extend radially across the flow passage. Circumferentially spaced-apart vanes also extend radially across the flow passage and define a plurality of inter-vane passages. Each of the struts is integrated to an associated one of the vanes to form therewith an integrated strut-vane airfoil. The inter-vane passages on either side of the integrated strut-vane airfoil may be adjusted for aerodynamic considerations. The vanes may be made separately from the struts and manufactured such as to cater for potential misalignments between the parts. | 09-11-2014 |
20140255147 | METHOD OF IMMOBILIZING LOW PRESSURE SPOOL AND LOCKING TOOL THEREFORE - A method of immobilizing a low pressure spool assembly including maintaining a body of the locking tool in the annular gas path, attaching a securing portion of the body across an aperture defined through an annular wall delimiting the gas path; positioning a stop connected to the body of the locking tool into a rotary path of a given one of the sets of blades of the low pressure spool assembly; and rotating the high pressure spool assembly thereby biasing a blade of the given set of blades of the low pressure spool assembly against the stop, thereby immobilizing the low pressure spool assembly. A locking tool and a method of performing engine maintenance are also provided. | 09-11-2014 |
20140252769 | GAS TURBINE ENGINE WITH INTERNAL ELECTROMECHANICAL DEVICE - A gas turbine engine including high and low pressure shafts, an electromechanical device having a rotor and a stator coupled such that the rotor is rotatable with respect to the stator, the rotor having a device gear secured thereto, the device being secured to a support structure in a bearing housing forming part of a bearing assembly supporting a portion of the low pressure shaft extending in proximity of the high pressure shaft and of the shaft gear, and a coupling idle gear secured for rotation about a stationary gear support mounted in the bearing housing, the idle gear being in toothed engagement with the shaft gear and with the device gear. An electromechanical device assembly for a gas turbine engine and a method of operating an electromechanical device are also provided. | 09-11-2014 |
20140252171 | NITROGEN BUBBLER SYSTEM IN FUEL TANK AND METHOD - An assembly of fuel tank and bubbler system comprises a fuel tank defining a storage volume adapted to receive fuel therein. The fuel tank comprises at least one fuel inlet, at least one fuel outlet, and at least one gas inlet adapted to be connected to an inert gas source to feed inert gas to fill an ullage in the storage volume. A bubbler system has at least one pipe extending into the fuel tank and adapted to be immersed into the fuel of the fuel tank. The pipe has a porous structure. The pipe is adapted to be connected to an inert gas source to inject inert gas via the porous structure into the fuel of the fuel tank. A method for treating fuel in a fuel tank is also provided. | 09-11-2014 |
20140251258 | SYSTEM FOR PILOT SUBCHAMBER TEMPERATURE CONTROL - There is described a system and method for controlling a temperature in the subchamber of a rotary engine. At least one first measurement of at least one engine operating parameter and a second measurement indicative of a present temperature in the subchamber are received. A setpoint for the temperature in the subchamber is computed from the at least one first measurement. In response to the second measurement, at least one control signal indicative of a request to adjust the present temperature towards the setpoint is generated and sent to the engine. | 09-11-2014 |
20140250897 | TIP-CONTROLLED INTEGRALLY BLADED ROTOR FOR GAS TURBINE ENGINE - An integrally bladed rotor for a gas turbine engine includes a hub, a plurality of blades radially extending from the hub and being integrally formed therewith. The hub having a rim from which the blades project and a pair of axially opposed split hub members extending at least radially inward from the rim. Each of the split hub members has a radially outer flex arm portion extending form the hub and a radially inner moment flange portion. At least one moment inducing element separately formed from the hub is mounted axially between the opposed split hub members and acts on the moment flange portions of the opposed split hub members to generate an inward bending moment on the flex arm portions of the opposed split hub members during rotation of the rotor, thereby deflecting the rim and the blades of the rotor radially inwardly. | 09-11-2014 |
20140250896 | COMBUSTOR HEAT SHIELD WITH CARBON AVOIDANCE FEATURE - The build-up of carbon deposition on the front face of a combustor heat shield is discouraged by jetting air out from the front face of the heat shield with sufficient momentum to push approaching fuel droplets or rich fuel-air mixture way from the heat shield. | 09-11-2014 |
20140250680 | METHOD OF ASSEMBLING AN ELECTROMECHANICAL DEVICE IN A GAS-TURBINE ENGINE - A method of assembling an electromechanical device in a gas-turbine engine, including mounting a rotor of the device on a rotor support, securing a stator of the device to a stator support, coupling the rotor support to the stator support such that said rotor is rotatable about said stator, securing the device to a bearing support, securing a bearing assembly on the low pressure shaft, coupling the device to the low pressure shaft by installing the bearing support over the bearing assembly, and drivingly engaging the rotor support to the high pressure shaft. | 09-11-2014 |
20140245988 | ROTARY INTERNAL COMBUSTION ENGINE WITH PILOT SUBCHAMBER - A rotary engine with an outer body having an insert located in the peripheral offset from the rotor cavity such that a portion of the peripheral wall extends between the insert and the cavity. The insert has a pilot subchamber defined therein and the portion of the peripheral wall has at least one opening defined therethrough in communication with the at least one outlet opening of the insert and with the cavity. A method of combusting fuel into a rotary engine is also discussed. | 09-04-2014 |
20140245745 | FUEL SYSTEM OF GAS TURBINE ENGINES - A method for purging fuel from a fuel system of a gas turbine engine on shutdown of the engine comprises, in one aspect, terminating a fuel supply to the fuel system and using the residual compressed air to create a reversed pressure differential in the fuel system relative to a forward pressure differential of the fuel system used to maintain fuel supply for engine operation, and under the reversed pressure differential substantially purging the fuel remaining in the system therefrom to a fuel source. | 09-04-2014 |
20140245744 | CONTROL OF GAS TURBINE ENGINE - Systems, devices, and methods for controlling a fuel supply for a turbine or other engine using direct and/or indirect indications of power output and optionally one or more secondary control parameters. | 09-04-2014 |
20140245739 | FUEL SYSTEM OF GAS TURBINE ENGINES - A method for purging fuel from a fuel system of a gas turbine engine on shutdown of the engine comprises, in one aspect, terminating a fuel supply to the fuel system and using the residual compressed air to create a reversed pressure differential in the fuel system relative to a forward pressure differential of the fuel system used to maintain fuel supply for engine operation, and under the reversed pressure differential substantially purging the fuel remaining in the system therefrom to a fuel source. | 09-04-2014 |
20140245581 | LOW INDUCTANCE CAPACITOR ASSEMBLY - A low-inductance capacitor assembly ( | 09-04-2014 |
20140244132 | SYSTEM AND METHOD FOR 0N-WING ENGINE TRIM VERIFICATION - Devices and methods relating to gas turbine engines and engine temperature trim verification are disclosed. An exemplary method comprises acquiring signals representing a plurality of engine parameters measured while the engine is operating and determining a recommended trim thermocouple resistance based at least partly on the measured parameters. The engine parameters may comprise at least an engine inlet temperature and an exhaust temperature. | 08-28-2014 |
20140241901 | IMPELLER - An impeller for increasing the pressure of a fluid circulating in an annular fluid path, the impeller comprising: a plurality of centrifugal compressor vanes circumferentially interspaced around the axis of the annular fluid path, the plurality of compressor vanes extending from an axially-oriented inlet to a radially-oriented outlet, and each having an inner edge and a free edge, the free edge of the plurality of compressor vanes coinciding with an outer limit of the annular fluid path, and a hub having a solid-of-revolution shape centered around an axis, the hub having an outer hub surface forming an inner limit to the annular fluid path and to which the inner edge of the plurality of centrifugal vanes is secured, the outer hub surface having a portion which leans forward, forming an axial recess therein. | 08-28-2014 |
20140241899 | BLADE LEADING EDGE TIP RIB - A rotor blade for a gas turbine engine includes a leading edge tip rib projecting outwardly from an airfoil of the blade at a tip region thereof. The tip rib continuously surrounds a leading edge of the airfoil and extends rearwardly from the leading edge along respective pressure and suction side surfaces to thereby alter the blade tip leakage vortex structure and strength, resulting in a stage efficiency benefit. | 08-28-2014 |
20140241863 | EXHAUST SECTION FOR BYPASS GAS TURBINE ENGINES - A turbine exhaust section comprises a turbine exhaust case having radially outer and inner ducts defining therebetween an annular exhaust portion for the hot exhaust gases, and an exhaust mixer projecting axially rearwardly from the turbine exhaust case for mixing the hot exhaust gases with a cooler bypass duct flow. The upstream end of the exhaust mixer surrounds a downstream end of the outer duct and defines therewith an axially extending overlap joint with a radial play between the outer duct and the exhaust mixer. A sliding attachment in the radial direction is provided between the outer duct and the exhaust mixer for accommodating differences in thermal growth during engine operation. The sliding attachment includes a circumferential array of sliding guides extending radially through the axially extending overlap joint. A resilient sealing ring seals the radial play between the outer duct and the exhaust mixer at the overlap joint. | 08-28-2014 |
20140241854 | ACTIVE TURBINE OR COMPRESSOR TIP CLEARANCE CONTROL - A gas turbine engine includes an annular plenum defined with an outer skin and a perforated inner skin for receiving selective air flow to impinge a support case which supports shrouds of the rotor assemblies of the engine therein for active tip clearance control of the rotor assemblies. In one embodiment a bobbin-type transfer tube for supplying cooling air into the plenum, is provided between an outer case of the engine an the plenum such that the thermally induced relative movement of the outer case and the plenum is permitted. | 08-28-2014 |
20140238642 | HEAT EXCHANGE DEVICE AND METHOD - A heat exchange device comprising a fluid flow passage having a plurality of successive segments in fluid flow communication with one another, the segments being adapted to maintain a developing flow therein and thereby improve heat transfer. | 08-28-2014 |
20140238038 | GAS TURBINE VARIABLE FOCUS LASER IGNITION - A laser ignition system for a gas turbine engine includes a combustion chamber. The system comprises a laser source for generating a continuous laser beam during an ignition process of the combustion chamber; and a dynamic laser focus apparatus positioned outside of the combustion chamber and focusing the laser beam into a continuously varying focal point to generate a laser kernel moving within a spray of air/fuel mixture injected into the combustion chamber. | 08-28-2014 |
20140237989 | LASER-IGNITION COMBUSTOR FOR GAS TURBINE ENGINE - The combustor has a laser ignitor mounted to the casing, remotely from the liner of the combustion chamber. The laser ignitor has an igniter beam path for igniting the fuel and air mixture in the combustion chamber, the igniter beam path extending at least partially across the air plenum surrounding the liner and into the combustion chamber through a corresponding beam path aperture provided in the liner. | 08-28-2014 |
20140237804 | ROTOR CENTRALIZATION FOR TURBINE ENGINE ASSEMBLY - One or more support elements radially extend through one or more openings defined in a turbine engine casing and are configured to centralize and at least partially support a rotor assembly of the engine during an engine disassembly or assembly procedure. The support elements are configured to transfer any rotor assembly weight loads to an engine casing while a bearing support of the rotor assembly is absent or removed. | 08-28-2014 |
20140234555 | METHOD OF PROTECTING A SURFACE - A method of masking part of a surface of a wall of a gas turbine component including at least one area having cooling holes defined therein, the method including applying a viscous curable masking compound to the part of the surface over an entirety of each of the at least one area, including blocking access to the cooling holes from the surface by applying the masking compound over the cooling holes without completely filling the cooling holes with the masking compound, and forming a respective solid masking element completely covering each of the at least one area and the cooling holes defined therein by curing the masking compound. | 08-21-2014 |
20140233284 | OVER VOLTAGE PROTECTION FOR ELECTRIC MACHINES - A system and method for protecting an electrical power generation system from an over-voltage. The output voltage of a multi-phase rectifier, operatively connected between the output terminals of an electric machine and a load, is monitored. The input of the multi-phase rectifier is short-circuited upon detection that the output voltage has reached a threshold voltage. Removal of the short-circuiting of the input of the multi-phase rectifier is synchronized with a substantially zero-crossing of phase current flowing through switching devices in the rectifier once the output voltage is no longer above the threshold voltage. | 08-21-2014 |
20140231247 | METAL PLATING METHOD AND APPARATUS - An apparatus and a method suited for metal plating aircraft engine components that allows the creation a local environment for plating by covering a localized area to be plated so that the localized area to be plated is sealed from remaining parts of the component, thereby eliminating the need for masking remaining parts of the component prior to plating. | 08-21-2014 |
20140231062 | LIQUID COOLING SYSTEM WITH THERMAL VALVE DEFLECTOR - The liquid cooling system has a heat exchanger having a fluid inlet and an outlet; a fluid supply conduit leading to the inlet of the heat exchanger; a fluid return conduit extending from the outlet of the heat exchanger; a bypass conduit extending between the fluid supply conduit and the fluid return conduit; a thermal valve configured for selectively closing the bypass conduit, the valve having a temperature sensing element positioned downstream of both the heat exchanger and the bypass conduit, the temperature sensing element configured to selectively move the thermal valve in response to a temperature change of the liquid which the temperature sensing element is exposed to relative to a temperature threshold of the valve; and a deflector positioned between the temperature sensing element and at least one of the bypass conduit and the heat exchanger outlet. | 08-21-2014 |
20140230404 | GAS TURBINE ENGINE EXHAUST MIXER - An exhaust mixer for a gas turbine engine includes an annular wall having upstream end adapted to be fastened to an engine case and a downstream end forming a plurality of inner and outer mixer lobes. A support member interconnects at least a number of the inner lobes, and includes a circumferentially extending stiffener ring located radially inwardly from the inner lobes and a series of circumferentially spaced apart mixer struts radially extending from the inner lobes to the stiffener ring. The mixer struts have a radial length at least equal to a width of a main gas path defined between the inner lobes and the exhaust cone such that the mixer struts extend entirely through the main gas path. The stiffener ring being fixed solely to the mixer struts such as to float with respect to the exhaust cone and permit relative movement therebetween. | 08-21-2014 |
20140219817 | HIGH PRESSURE TURBINE BLADE COOLING HOLE DISTRIBUTION - A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3. | 08-07-2014 |
20140219816 | HIGH PRESSURE TURBINE BLADE COOLING HOLE DISTRIBUTION - A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3. | 08-07-2014 |
20140208760 | GAS TURBINE ENGINE WITH TRANSMISSION - A gas turbine engine with a transmission having a first rotatable member coupled to an engine spool, a second rotatable member coupled to a compressor rotor, and coupled rotatable members defining at least first and second alternate transmission paths between the first and second members. Each transmission path defines a different fixed transmission ratio of a rotational speed of the second member on a rotational speed of the first member. | 07-31-2014 |
20140173899 | GAS TURBINE ROTOR ASSEMBLY METHODS - Method of assembling a rotor assembly of a gas turbine engine having a plurality of components. The method comprises in one aspect calculating the bending forces due to the mass distribution along the rotor. In another aspect, an optimization routine iterates different rotor arrangements, comparing the calculated bending moments to determine a set of component positions that minimizes the bending forces. In another aspect, mass corrections are optimized to balance the rotor assembly. | 06-26-2014 |
20140165533 | GAS TURBINE ENGINE MOUNTING RING - A casing for an aircraft engine includes an outer ring and an inner hub defining an airflow passage therebetween, the outer ring having an axis defining an axial direction; a plurality of struts arranged in a circumferential array and extending radially from the inner hub to the outer ring to mount the inner hub to the outer ring; wherein the outer ring is defined by a double skin including an axially-extending annular outer skin of sheet metal concentrically surrounding and radially-spaced from an annular inner skin of sheet metal, the outer and inner skins generally parallel to one another, an annular front end ring and an annular rear end ring welded or brazed to the outer and inner skins adjacent respective front and rear edges of the skins to define an annular cavity between them, and the outer ring further comprising a plurality of circumferentially spaced axially-extending ribs interconnecting the outer and inner skins to reinforce the double skins. | 06-19-2014 |
20140154053 | CARBON SEAL ASSEMBLY - A carbon seal assembly comprises an annular member adapted to be secured to a structure. An annular seal runner is sealingly mounted to a shaft to rotate therewith, with the seal runner being made of a material complementary to that of the annular member for magnetic attraction therebetween. An annular carbon seal element is mounted to the annular seal runner to rotate therewith and positioned in a gap between the annular member and the annular seal runner, the annular carbon seal element having an annular wear surface abutting against a face of the annular member. A cross-sectional area of the annular carbon seal element increases as the axial dimension of the annular seal runner decreases for at least a part of the seal. | 06-05-2014 |
20140150550 | METHOD AND SYSTEM FOR INTEGRATING GAS TURBINE TRIM BALANCING SYSTEM INTO ELECTRONIC ENGINE CONTROLS - Methods and systems for determining a rebalancing strategy for trim balancing one or more rotational components of a gas turbine. One or more noise, acoustics or vibrational signals may be received at a control device of an aircraft comprising the gas turbine engine while the aircraft is in operation. The one or more noise, acoustics or vibrational signals may be used for determining a rebalancing strategy for one or more unbalanced rotational components. | 06-05-2014 |
20140150449 | AIR COOLING SHAFT AT BEARING INTERFACE - A gas turbine engine including a compressor rotor and a turbine rotor connected by a compressor shaft portion connected to the compressor rotor and a turbine shaft portion connected to the turbine rotor. The compressor shaft portion and the turbine shaft portion are connected axially together by a shaft coupling, between the compressor rotor and the turbine rotor, and at least a bearing rotatably coupled to the compressor shaft portion adjacent the shaft coupling. The compressor shaft and/or the turbine shaft are provided with openings permitting cooling air to enter air passages in the area of the shaft coupling and surrounding the end of the turbine shaft portion, in order to dissipate heat originating at the turbine rotor and thus reducing the thermal stresses at the bearing. | 06-05-2014 |
20140148942 | METHODS AND SYSTEMS FOR CALCULATION OF FEEDRATE - Methods and systems for calculating a feedrate for programming a multi-axis machining tool. For at least one control block in a defined machining path: a displacement of a defined machine control point from a previous control block to a current control block is determined; a displacement of a defined feedrate control point from the previous control block to the current control block is determined; a compensation ratio is calculated as a ratio between the displacement of the defined machine control point and the displacement of the defined feedrate control point; and a feedrate for the machine control point is calculated by applying the compensation ratio to a desired feedrate. The calculated federate is used in a control block of a multi-axis machining tool. | 05-29-2014 |
20140147320 | WANKEL ENGINE ROTOR - A rotor for a Wankel engine including a plurality of ribs extending from a bearing support to each one of the flanks, the plurality of ribs including, for each flank, first and second ribs connected to the flank between the recess and a respective one of the apex portions. The first and second ribs are curved along at least a portion thereof, and/or the first and second ribs are closest to the respective apex portion and connected to the flank adjacent a junction between a portion of the flank defining the recess and a respective portion of the flank connected thereto. A method of reducing pinching of apex seals in a rotor of a Wankel engine is also discussed. | 05-29-2014 |
20140145525 | REDUCED COGGING TORQUE PERMANENT MAGNET MACHINE - An electric machine is formed by a stator and a rotor that is free to rotate about an axis of rotation. The stator may have teeth projecting from a body portion and that define slots for housing electrical windings. The rotor may have a rotor core and a number of magnets supported on a peripheral face of the rotor in substantially contiguous arrangement and of alternating magnetization. The rotor magnets are shaped so that pairs of adjacent magnets oppose one another along magnetic boundary lines that are skewed relative to the slots formed in the body portion of the stator. For example, the shape of the rotor magnets may be arcuate trapezoidal or parallelogramatic. In this configuration, cogging torque experienced by the rotor during operation of the electric machine may be reduced. | 05-29-2014 |
20140144671 | MULTI-PHASE CABLE - A multi-phase cable, the cable including a plurality of conductors for conducting currents of two or more different phases, each phase being associated with one or more conductors and each conductor being associated with one respective phase. Each conductor has a cross-section with at least one dimension that is sized to decrease a skin effect of the conductor at a maximum or nominal operation frequency of the conductor. The conductors are arranged to permit free air cooling of the cable on at least two sides of each conductor, and such that each conductor of a given phase has, as immediate neighbors, only conductors of one or more different phases. | 05-29-2014 |
20140144154 | GAS TURBINE ENGINE WITH BEARING BUFFER AIR FLOW AND METHOD - The gas turbine engine has a bleed air aperture formed in the radially outer wall upstream from the combustor and a bearing cavity formed within the radially inner wall, at least two bearing seals enclosing at least one bearing in the bearing cavity and separating the bearing cavity from associated buffer air entry points, an oil supply system including oil paths leading to each of the bearings; a buffer air supply system including buffer air paths leading to each of the entry points. | 05-29-2014 |
20140144139 | AIR COOLED AIR COOLER FOR GAS TURBINE ENGINE AIR SYSTEM - An air-to-air cooler has a heat exchanger integrated to a housing. A first passage extends through the housing for directing a flow of cooling air through the heat exchanger. A second passage extends through the housing for directing a flow of hot air to be cooled through the heat exchanger. The first passage has a cooling air outlet tube disposed downstream of the heat exchanger. The cooling air outlet tube extends across the second passage between the heat exchanger and a hot air inlet of the second passage. The hot air inlet is disposed to cause incoming hot air to flow over the cooling air outlet tube upstream of the heat exchanger. An ejector drives the flow of cooling air through the first passage of the air-to-air cooler. A portion of the hot air flow may be used to drive the ejector. | 05-29-2014 |
20140144121 | GAS TURBINE ENGINE WITH BEARING OIL LEAK RECUPERATION SYSTEM - A gas turbine engine having an annular gas path between a radially outer wall and a radially inner wall, leading successively across at least one compressor stage, a combustor section, and at least one turbine stage, a hollow shaft having an internal surface with an oil trap formed therein, and at least one oil recuperation orifice extending out across the hollow shaft from the oil trap; and a bearing cavity formed within the radially inner wall, having at least one bearing therein rotatably supporting the hollow shaft of the gas turbine engine, at least two bearing seals enclosing the at least one bearing in the bearing cavity and separating the bearing cavity from associated buffer air entry points, at least a first one of said buffer air entry points being exposed to the at least one oil recuperation orifice outside the hollow shaft. | 05-29-2014 |
20140133962 | INTERTURBINE VANE WITH MULTIPLE AIR CHAMBERS - A gas turbine engine has a mid turbine frame disposed between turbine rotor assemblies. The mid turbine frame includes hollow airfoils radially extending through an annular gas path duct. The airfoils each include a double-walled leading edge structure to define a front chamber separated from a rear chamber defined in the remaining space within the airfoil. | 05-15-2014 |
20140123663 | ROTOR RESONANCE DISTURBANCE REJECTION CONTROLLER - A speed control system for an engine comprising at least one rotary load is provided. The speed control system may include a rotor speed controller configured to regulate speed in the rotary load based on a sensed rotor speed, exclusive of resonant mode speed oscillations, in closed loop feedback with a commanded rotor speed. To provide active damping of resonant mode speed oscillations, a resonance disturbance rejection controller may be configured to compensate a speed control signal by observing a component of the sensed rotor speed that is due to resonant mode oscillations. Based on the observed resonance component, the resonance disturbance rejection controller may compute an adjustment value for the speed control signal. In the particular case of gas turbine engines, the resonance disturbance rejection controller may effect active damping by compensation of a fuel flow request for a gas generator. | 05-08-2014 |
20140121994 | METHOD AND SYSTEM FOR FAILURE PREDICTION USING LUBRICATING FLUID ANALYSIS - Methods and systems for failure prediction using analysis of oil or other lubricant. Raw data about feature(s) of each of a plurality of particles filtered from a fluid sample are used to categorize each particle into one of a plurality of categories, each category being defined by one or more of: chemical composition, size and morphology. Particle physical characteristics in each category are quantified to obtain a set of categorized data. The categorized data are compared with historical data. Results of the comparing are evaluated to generate a prediction of any failure or mechanism of failure. | 05-01-2014 |
20140121861 | CONFIGURABLE AIRCRAFT DATA ACQUISITION AND/OR TRANSMISSION SYSTEM - An aircraft data acquisition and/or transmission system comprises a memory having stored therein a first state machine configuration comprising a first allocation between a plurality of transition criteria objects stored in a first area of the memory and a plurality of state objects stored in a second area of the memory separate from the first area. A processor is coupled to the memory and at least one application executable by the processor causes a first data acquisition and/or transmission behavior in the aircraft by running the first state machine configuration, receives a second state machine configuration comprising a second allocation between the plurality of transition criteria objects and the plurality of state objects, and causes a second data acquisition and/or transmission behavior in the aircraft by running the second state machine configuration. | 05-01-2014 |
20140116061 | COUPLING ELEMENT FOR TORQUE TRANSMISSION IN A GAS TURBINE ENGINE - A gas turbine engine having a coupling element for coupling a first shaft to a second shaft, the second shaft being substantially axially aligned with the first shaft, the coupling element provided with an exterior surface that engages an opposing interior surface of the first shaft and an interior surface that engages an opposing exterior surface of the second shaft to facilitate torque transfer between the first shaft and the second shaft when rotated together, wherein the coupling element and at least one of the first and second shafts cooperate to define at least one fluid passageway therebetween. | 05-01-2014 |
20140103884 | RESONANT MODE DAMPING SYSTEM AND METHOD - A control system for increasing the damping of a resonant mode of a rotor drive train is provided. A control loop is introduced in parallel with the free turbine speed governing control loop. The control loop receives a feedback signal from the free turbine and rotor drive train, the feedback signal representative of the free turbine speed. The control loop extracts the resonant component from the feedback signal to generate a resonant signal containing the resonant component and estimates a derivative of the free turbine speed from the resonant signal. The control loop then outputs a request for damping the resonant mode on the basis of the derivative. A narrow-band reject filter is introduced in the free turbine speed governing loop to filter out the resonant component from the feedback signal, thereby ensuring that the free turbine speed governing loop does not interfere with the control loop. | 04-17-2014 |
20140101939 | METHOD OF FABRICATING INTEGRALLY BLADED ROTOR AND STATOR VANE ASSEMBLY - A method of fabricating an integrally bladed rotor of a gas turbine engine according to one aspect, includes a 3-dimensional scanning process to generate a 3-dimensional profile of individual blades before being welded to the disc of the rotor. A blade distribution pattern on the disc is then determined in a computing process using data of the 3-dimensional profile of the individual blades such that the fabricated integrally bladed rotor is balanced. | 04-17-2014 |
20140095051 | ADAPTIVE FUEL MANIFOLD FILLING FUNCTION FOR IMPROVED ENGINE START - There is described a system and method for filling an engine fuel manifold. An adaptive filling function is used to determine a flow rate at which fuel is to be delivered to the fuel manifold. The filling function receives as an input a present measurement of the engine's speed and computes the flow rate accordingly. The fuel manifold may then be filled according to the computed flow rate so as to match the engine's speed. Appropriate fuel/air ration conditions can therefore be achieved for successful engine start. | 04-03-2014 |
20140086756 | INTERNALLY COOLED GAS TURBINE ENGINE AIRFOIL - A gas turbine engine airfoil has a hollow airfoil section extending chordwise between a leading edge and a trailing edge. The airfoil has a leading edge cooling passage and a separate serpentine passage for cooling a remaining portion of the airfoil. The serpentine passage has at least three segment serially interconnected in fluid flow communication. The leading edge cooling passage and the serpentine cooling passage have separate coolant inlets. The coolant inlet of the serpentine passage comprises a primary inlet branch connected in fluid flow communication with a first one of the segments of the serpentine passage and a secondary inlet branch connected in flow communication with a last one of the segments, thereby providing for a portion of the flow passing through the coolant inlet of the serpentine passage to be directly fed into the last segment of the serpentine passage. | 03-27-2014 |
20140084130 | GEARBOX POSITIONING DEVICE - A gearbox positioning device to position a gearbox relative to an engine case of a gas turbine engine. The positioning device has a first and second part, with the first part having a connecting cavity of predetermined shape and being secured to the engine case. The second part has a coupling post received in the connecting cavity and constraining displacement of the gearbox in a laterally transverse direction with respect to the central longitudinal axis of the engine. The device permits movement of the gearbox in the axial direction of the engine as well as angular movement to prevent stressing the engine case during thermal growth of the engine case or support linkage imbalance. A gearbox positioning system incorporating the positioning device and the method of attaching same to the engine casing is also included. | 03-27-2014 |
20140079530 | AIR COOLING DESIGN FOR TAIL-CONE GENERATOR INSTALLATION - A system for cooling a generator mounted in the tail-cone of an engine. The system comprises a fairing, which receives through an inlet thereof air from a bypass duct and directs the bypass air towards a cavity of the tail-cone for cooling the generator. The bypass air is then expelled through an outlet of a support strut positioned in fluid communication with the tail-cone cavity. The fairing inlet and the strut outlet are both positioned in a plane substantially perpendicular to a longitudinal plane of the engine. In this manner, circulation of the bypass air through the fairing, the tail-cone cavity, and the strut may be achieved. The bypass air directed through the fairing further enables cooling of service lines accommodated in the fairing. A lobe mixer is further used to direct the fairing and shield the latter from core exhaust. | 03-20-2014 |
20140077507 | PRESSURIZATION OF GENERATOR - Systems and methods for pressurization of a generator in a gas turbine engine. The system may include a fluid valve for permitting or inhibiting airflow into a housing of the generator. The fluid valve may be in fluid communication with an air source within the engine, and may be configured to permit airflow into the housing of the generator when pressure within the housing is below a preset pressure threshold or range. | 03-20-2014 |
20140075919 | TEC MIXER WITH VARIABLE THICKNESSES - A mixer of a bypass turbine aeroengine according to one embodiment, includes circumferential inner and outer flow surfaces in a wavy configuration to form a plurality of lobes of the mixer. The mixer has an upstream end portion of sheet metal with a first thickness and a downstream end portion of sheet metal with a second thickness less than the first thickness. | 03-20-2014 |
20140073225 | CHANNEL INLET EDGE DEBURRING FOR GAS DIFFUSER CASES - A method of deburring channel inlet edges inside a cavity of a gas diffuser case is disclosed. The diffuser case has a plurality of channels each having an inner surface and an inlet edge defining an inlet of the channel. The surfaces of adjacent channels co-operate to provide said inlet edge therebetween. The inlet edges of the channels are provided in an inwardly facing circular array around a central axis of the gas diffuser case. The method comprises: inserting a tool head having at least one nozzle in the cavity of the gas diffuser case; and then ejecting abrasive particles from at least one nozzle towards at least one of the channel inlet edges of the gas diffuser case to at least one of decrease a radius of at least one said edge and improve a smoothness of at least one said surface. | 03-13-2014 |
20140069105 | COMPRESSOR SURGE PREVENTION DIGITAL SYSTEM - Methods and devices for anticipating a surge in a gas turbine engine. Controlled pressure signal(s) may be compared with reference pressure signal(s), each of the controlled pressure signal(s) and reference pressure signal(s) having an associated time value. If the controlled pressure signal(s) are less than the reference pressure signal(s), a controlled pressure curve may be fitted through a predetermined number of points based on the controlled pressure value(s) and associated time value(s). A reference pressure curve may be fitted through the predetermined number of points based on the reference pressure value(s) and associated time value(s). A time to compressor surge may be estimated based on an intersection of the controlled pressure curve and the reference pressure curve. | 03-13-2014 |
20140060074 | SYSTEMS AND METHODS FOR DRIVING AN OIL COOLING FAN OF A GAS TURBINE ENGINE - Systems and methods for driving an oil cooling fan ( | 03-06-2014 |
20140055103 | SYSTEM AND METHOD FOR MONITORING TEMPERATURE INSIDE ELECTRIC MACHINES - The electric machine comprises at least one winding made of a material having a temperature dependent resistance. The temperature of the winding is monitored using the resistance therein. Temperatures or resistances indicative of a fault can be sensed, and corrective action taken, without the need for dedicated temperature sensors. | 02-27-2014 |
20140046478 | METHOD OF MAKING A PART AND RELATED SYSTEM - A system and method of making a part, including capturing an actual tridimensional surface of each part to obtain a corresponding digitized actual surface, performing a tridimensional comparison between the digitized actual surface of each part and a nominal tridimensional surface, generating an actual location and orientation of an operation to be performed based on the tridimensional comparison, and performing the operation on the part based on the generated actual location and orientation of the element. | 02-13-2014 |
20140038496 | ACTIVE COOLANT FLOW CONTROL FOR MACHINING PROCESSES - A method for operating a machining tool, comprising: setting a flow rate and a pressure of a flow of coolant to a target flow rate and a pressure target, respectively, the coolant flow being provided to the machining tool; machining a work-piece using the machining tool; measuring the flow rate and the pressure of the flow of coolant; and detecting an anomaly with respect to the coolant flow; and taking a corrective action depending on the type of the detected anomaly. | 02-06-2014 |
20140037396 | APPARATUS FOR PROVIDING FIRTREE SLOTS - A apparatus for forming a firtree slot in a disc of a bladed rotor assembly for a gas turbine engine may include a tip portion having the shape of the profile of one side of the firtree slot to be formed but smaller than the profile of the complete firtree slot so that the tool will only engage one side of the firtree slot. | 02-06-2014 |
20140033730 | PRESSURE REGULATING VALVE FOR AIRCRAFT ENGINE - The fluid pressure regulating valve is for use in an aircraft engine. The valve has at least a first fluid inlet, a first fluid outlet and a second fluid outlet. The valve comprises: a valve housing having a first valve spool interior cavity; a first fluid path within the valve housing from the fluid inlet to the first fluid outlet; a second fluid path within the valve housing from the fluid inlet to the second fluid outlet; a first valve spool mounted for reciprocal motion within the first valve spool cavity between a first position and a second position, the first valve spool having a second valve spool interior cavity and being spring-biased to its first position; and a second valve spool mounted for reciprocal motion within the second valve spool cavity between a first position and a second position, the second valve spool being spring-biased to its first position, the second valve spool closing the first fluid path at its second position when the first valve spool is substantially at its first position. | 02-06-2014 |
20140026591 | AXIAL RETENTION FOR FASTENERS IN FAN JOINT - In a gas turbine engine, a fan rotor and a compressor rotor are connected to a joint which is attached to a shaft. The fan rotor is connected to the joint by a plurality of fasteners extending through mounting openings in the joint and through apertures in the fan rotor. A standard retaining ring is attached to the joint to cover a portion of an enlarged head of each of the respective fasteners in order to prevent removal of the respective fasteners from the mounting openings in the joint before the respective fasteners are connected to the fan rotor. | 01-30-2014 |
20140026415 | GAS TURBINE ROTOR ASSEMBLY METHOD - Methods of assembling a rotor assembly of a gas turbine engine having a plurality of components. The method comprises, in one aspect, determining a static unbalance force of each component using the mass and the center of mass location, and determining an optimum assembly arrangement minimizing static unbalance forces. Optimal component stacking positions, among other things, may then be established. | 01-30-2014 |
20140021958 | SOLENOID TRANSIENT VARIABLE RESISTANCE FEEDBACK FOR EFFECTER POSITION DETECTION - Feedback from a solenoid is achieved by adding at least one variable resistance in parallel with the solenoid current feedback circuit for position detection. The resistance has current flowing therethrough when a switching device actuated by the solenoid is in one position or transitions from one position to at least one other position. A feedback current may be measured in the current feedback circuit and the position of the switching device in response to actuation thereof by the solenoid may be determined from the measured feedback current. | 01-23-2014 |
20140017477 | Thermal Barrier Coating with Lower Thermal Conductivity - A thermal barrier coating includes a microstructure and a composition including: a ceramic based compound comprising gadolinia and zirconia. The coating includes a nano-structure having a porosity of at most 50% by volume of the coating. | 01-16-2014 |
20140012481 | AIRCRAFT ENGINE CONTROL DURING ICING OF TEMPERATURE PROBE - Methods for controlling an aircraft turbofan engine during icing of a temperature probe and devices for carrying out such methods are described. The methods may comprise: using one or more signals representative of temperature received from a heated temperature probe to generate one or more control signals for use in controlling the engine; determining that an icing condition associated with the probe exists; and using data representing one or more substitute signals in place of signals representative of temperature received from the heated temperature probe to generate the one or more control signals for use in controlling the engine. | 01-09-2014 |
20130291554 | AIR COOLER SYSTEM FOR GAS TURBINE ENGINES - A buffer air cooler system for gas turbine engines disposed in a bypass duct of the engine, includes a housing for containing the buffer air cooler therein and an inlet portion attached to the housing. In one embodiment, the inlet portion has a double-skin configuration in at least one region of a top, bottom and sides of the inlet portion. | 11-07-2013 |
20130288498 | CONNECTOR FOR MULTI-PHASE CONDUCTORS - A connector for multi-phase conductors. The connector includes conductors for conducting multi-phase currents, and at least one conductive plate corresponding to each phase. Each conductive plate defines apertures for the conductors to pass through, there being at least one aperture in each conductive plate for each respective conductor. Each conductive plate includes at least one connecting member for forming an external electrical connection. Each conductor passes through a respective aperture of each conductive plate. Each conductor is selectively coupled to form electrical connection only to a conductive plate corresponding to the respective phase of the conductor. | 10-31-2013 |
20130263603 | NOZZLE DESIGN TO REDUCE FRETTING - A method of designing a fuel nozzle of a gas turbine engine to reduce fretting thereof during use, including establishing an initial nozzle design, determining a first natural frequency of that design and a running frequency range of the gas turbine engine, and increasing a first transverse dimension of the stem member of the nozzle across a length of a portion thereof adjacent the inlet end until the first natural frequency of the nozzle is outside the running range, while a second transverse dimension of the portion remains at least substantially unchanged across the length thereof. | 10-10-2013 |
20130255268 | TURBINE ENGINE HEAT RECUPERATOR SYSTEM - A gas turbine engine recuperator system includes a heat recuperator positioned in a gas turbine exhaust gas duct for recovering heat from turbine exhaust gases to preheat a compressor flow being supplied to the combustor. A continuous bleed flow of the turbine exhaust gases is provided to bypass the heat recuperator. The continuous bleed flow of the turbine exhaust gases is adjustable to reduce turbine exhaust gas pressure loss at a high engine operation level and to provide efficient heat recovery at low and/or medium engine operation levels. | 10-03-2013 |
20130249338 | TANDEM ELECTRIC MACHINE ARRANGEMENT - The tandem electric machine arrangement comprises an outside rotor having two axially spaced-apart sets of circumferentially-disposed permanent magnets. It also comprises an inside stator having at least two electrically-independent windings, the at least two windings axially spaced apart from one another and disposed relative to the magnet sets to thereby be magnet coupled to a respective one of the sets of permanent magnets during rotation of the rotor. One of the rotor and the stator is provided in two separate pieces, each piece supported from opposite axial sides of the electric machine relative to one another. The other of the rotor and the stator is supported from substantially centrally of the two pieces. | 09-26-2013 |
20130202442 | FAN AND BOOST JOINT - A gas turbine engine having at least one spool assembly, the at least one spool assembly including a fan rotor, a compressor disposed downstream of the fan rotor, a turbine and a shaft connecting the fan rotor, compressor and turbine, a joint affixed to an upstream end of the shaft, and including a first link connecting the fan rotor to the shaft and a second link connecting the compressor to the shaft, the second link being less rigid then the first link. | 08-08-2013 |
20130202440 | SEGMENTED RINGS WITH CAPTIVE NUTS FOR FAN BOLTS - An apparatus and method for connecting a fan rotor to a shaft of a gas turbine engine includes a segmented clinch nut plate having a plurality of clinch nuts installed in pre-drilled holes in each segment. A plurality of mounting bolts extending through the holes in the rotor shaft and holes in a mounting device of the shaft threadedly engage the clinch nuts of the clinch nut plate. The segments of the clinch nut plate are disposed separately one from another. | 08-08-2013 |
20130199152 | TURBINE ENGINE HEAT RECUPERATOR PLATE AND PLATE STACK - A heat recuperator includes a plurality of channel walls composed substantially of thermally-conductive material and supported in spaced-apart relation, defining fluid channels and interstices therebetween. The fluid channels receive at least one primary fluid flow and the interstices receive at least one secondary fluid flow so as to effect heat exchange between the two flows. In use, the plurality of channel walls are deformable by pressure differential between the primary and secondary fluid flows. When at least some of the channel walls are in a deformed state, the plurality of channel walls are stabilized through press fit engagement of mutually opposed contact regions formed in adjacent pairs of the channel walls. | 08-08-2013 |
20130189071 | OIL PURGE SYSTEM FOR A MID TURBINE FRAME - An oil purge system for a mid turbine frame (MTF) of a gas turbine engine has an oil transfer tube surrounded by a heat shield tube. The oil transfer and heat shield tubes extend at their respective inner ends downwardly from an oil port of a bearing housing and terminate at their respective outer ends projecting outwardly from an annular wall of an outer case of the MTF. Oil leaked from the oil port is purged by pressurized air through an annular cavity formed between the oil transfer and heat shield tubes, and is discharged out of the MTF. | 07-25-2013 |
20130177410 | CASING FOR AN AIRCRAFT TURBOFAN BYPASS ENGINE - A casing for an aircraft turbofan bypass engine includes an outer ring, an inner hub and a plurality of struts radially extending therebetween. An annular portion of an engine core casing having an outer wall and an inner wall, is disposed between the outer ring and inner hub, forming an annular splitter supporting an upstream splitter tip structure. The annular splitter further includes an intermediate wall disposed in the annular splitter fixed to the outer wall and the struts, to distribute loads from the annular splitter box to the struts. | 07-11-2013 |
20130175327 | STRUCTURAL REINFORCEMENT STRUT FOR GAS TURBINE CASE - A gas turbine engine case having a working fluid flow, includes inner and outer case portions defining an annular duct for directing the working fluid flow, and a plurality of struts positioned within the annular duct and extending between the inner and outer case portions. The struts are welded to the inner and outer case portions with a first weld along a peripheral line of the respective struts and with second weld, of fillet type, in selected locations for additionally connecting a portion of each strut to the respective inner and outer case portions. | 07-11-2013 |
20130156589 | TURBINE ROTOR RETAINING SYSTEM - A rotor retaining system for a rotor assembly of a gas turbine engine includes a washer interlocking multiple retaining nuts to prevent relative rotation. The retaining nuts retain a rotor disc and a cover plate in position, respectively. A wire is provided to retain the washer in position and which may dampen vibration during engine operation. | 06-20-2013 |
20130156558 | ANNULAR GAS TURBINE ENGINE CASE AND METHOD OF MANUFACTURING - The method is used for making an annular gas turbine engine case from a preform. The method comprises comprising flowforming at least one section of the preform, and then outwardly bending at least one portion of the perform. | 06-20-2013 |
20130156541 | ACTIVE TURBINE TIP CLEARANCE CONTROL SYSTEM - An active tip clearance control (ATCC) system of a gas turbine engine includes an ejector to selectively drive an air flow passing through the ATCC system. A high pressure air flow as a motive flow of the ejector is controlled by a valve according to engine operation requirements. | 06-20-2013 |
20130139513 | FUEL NOZZLE AND METHOD OF REPAIR - A modular fuel nozzle tip for a gas turbine engine includes a body defining one or more fuel conveying passages extending therethrough, an annular cap having a radially inner shoulder surface interfacing with the peripheral end surface of the body to define a plurality of air channels extending through the head portion of the modular fuel nozzle tip. At least two fasteners fasten the annular cap to the body. | 06-06-2013 |
20130133332 | SYSTEMS AND METHODS FOR CHANGING A SPEED OF A COMPRESSOR BOOST STAGE IN A GAS TURBINE - Systems and methods for changing a speed ratio between a compressor boost stage ( | 05-30-2013 |
20130089416 | FABRICATED GAS TURBINE DUCT - A vane structure of a gas turbine engine includes a plurality of vanes extending radially between outer and inner shrouds. At least the outer shroud is formed substantially from a single-piece annular skin of sheet metal. A plurality of reinforcing plates are placed against and are affixed to an outer surface of the skin of the outer shroud in respective joining locations where a radial outer end of each of the respective vanes joins the skin. | 04-11-2013 |
20130086919 | STARTING OF AIRCRAFT ENGINE - A multi-engine system | 04-11-2013 |
20130086805 | ROTOR DISC AND METHOD OF BALANCING - A rotor disc, such as one made of a damage intolerant material or other material sensitive to stress concentrations, has at least one balancing assembly which includes a plurality of circumferentially spaced-apart sacrificial protrusions projecting between adjacent stress-relieving slots. Selective material removal is permitted from the rotor disc, while managing stress concentrations in the rotor disc created by such material removal, such that the rotor disc may be balanced without detrimentally affecting its service life. | 04-11-2013 |
20130074334 | ROTOR CENTRALIZATION FOR TURBINE ENGINE ASSEMBLY - One or more support elements radially extend through one or more openings defined in a turbine engine casing and are configured to centralize and at least partially support a rotor assembly of the engine during an engine disassembly or assembly procedure. The support elements are configured to transfer any rotor assembly weight loads to an engine casing while a bearing support of the rotor assembly is absent or removed. | 03-28-2013 |
20130071245 | ROTOR WITH IMPROVED BALANCING FEATURES - A rotor assembly for a gas turbine engine including a circumferential array of regularly spaced apart features provided on the disk which are each configured for receiving a balancing weight. The circumferential array of features has an angular orientation relative to the circumferential array of blades such that the features are each located at an at least substantially same offset angle from a stacking line of a respective adjacent blade. | 03-21-2013 |
20130067930 | AXIAL BOLTING ARRANGEMENT FOR MID TURBINE FRAME - A mid turbine frame of a gas turbine engine has a plurality of circumferentially spaced load transfer spokes extending radially between outer and inner cases. The load transfer spokes have a circumferentially enlarged inner end which is mounted to the inner case by a plurality of axially disposed fasteners extending through the inner case and spokes. | 03-21-2013 |