Abstract: A thermal management system for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a heat exchanger and a valve that controls an amount of a first fluid that is communicated through the heat exchanger A first sensor senses a first characteristic of a second fluid that is communicated through the heat exchanger to exchange heat with the first fluid and a second sensor senses a second characteristic of the second fluid. A positioning of the valve is based on at least one of the first characteristic and the second characteristic.

Abstract: A gas turbine engine includes a fan, a fan drive gear system coupled to drive the fan about an engine central axis, a compressor section including a first compressor and a second compressor and a turbine section. The turbine section includes a first turbine coupled to drive a first spool. The first spool is coupled at a first axial position to a compressor hub that is coupled to drive the first compressor. The first spool is also coupled at a second, different axial position to a fan drive input shaft that is coupled to drive the fan drive gear system. The turbine section also includes a second turbine coupled through a second spool to drive the second compressor. A sensor probe is operable to determine a rotational speed of the first spool. The sensor probe is located at a third axial position that is axially forward of the first axial position and axially aft of the second axial position.

Abstract: Apparatus and methods are disclosed for uniformly mixing fluid phases entrained in a porous material. A mixer may have a vessel and at least one porous material held by the vessel. At least one actuator may be acoustically coupled with at least one wall of the vessel for generating a wave. The wave effects mixing of at least two fluids in the porous material. The actuator may be a linear motor actuated with a control signal of predetermined frequency. The actuator may have a number of actuator pairs each including respective first and second actuators at respective first and second sides of the vessel. The actuators may be hinged for reciprocal movement. The actuators may be actuated to form a compression expansion wave to effect fluid motion in the porous material.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: A gas turbine engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section includes a fan drive turbine and a second turbine. The second turbine is disposed forward of the fan drive turbine. The fan drive turbine includes at least three rotors and at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between about 2.00 and about 2.30. A speed change system is driven by the fan drive turbine for rotating the fan about the axis.

Abstract: A turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, a high pressure turbine in fluid communication with the combustor, a low pressure turbine in fluid communication with the high pressure turbine, and a mid turbine frame located axially between the high pressure turbine and the low pressure turbine. The mid turbine frame includes an outer frame case, an inner frame case, and a plurality of hollow spokes that distribute loads from the inner frame case to the outer frame case. The spokes are hollow to allow cooling airflow to be supplied through the spokes to the inner frame case.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft.

Abstract: An aluminum alloy component has a surface region alloyed with an anodic metal to increase corrosion resistance in aqueous environments with high salinity or sulfur content.

Abstract: A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine.

Abstract: An example diffusion bonding machine includes a support structure configured to receive first and second die sets. A heat transfer device is arranged near the support structure and is configured to transfer heat relative to the die sets. A mechanism is configured to separate the die sets from one another during heat transfer. In one example method of diffusion bonding, heat is transferred relative to a space between die sets. The die sets are supported on the support structure, and a load is applied to the die sets to diffusion bond a component within each of the die sets.

Abstract: A method of improving the aerodynamic performance of an inlet section of a nacelle assembly of a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, (a) sensing an operability condition, (b) adjusting a leading edge of the inlet section in response to the step (a), and (c) adjusting a thickness of the inlet section in response to the step (a).

Abstract: A gas turbine engine includes a central body support that provides an inner annular wall for a core flow path. The central body support includes a first mount feature. A geared architecture interconnects a spool and a fan rotatable about an axis. A flex support interconnects the geared architecture to the central body support. The flex support includes a second mount feature that cooperates with the first mount feature for transferring torque there between. A method of disassembling a front architecture of a gas turbine engine includes the steps of accessing forward-facing fasteners that secure a central body support to a flex support, wherein the flex support includes a geared architecture supported thereon, removing the fasteners, and decoupling first and second mount features respectively provided on the central body support and the flex support.

Abstract: An integrally bladed disk includes a rotor disk and circumferentially spaced first and second blades. The rotor disk has a rim the periphery of which forms a flow surface. The first and second blades extend integrally outward from the rim. The rim defines a trench in the flow surface between the first and second blades aft of a leading edge of the rim. The trench extends axially forward and rearward of a leading edge of the first blade.

Abstract: A method for controlling a temperature of a portion of a die casting system having a shot tube plunger, according to an exemplary aspect of the present disclosure includes, among other things, communicating a fluid through a fluid inlet of a fluid passageway of a thermal control scheme of the shot tube plunger. The fluid circulates through the fluid passageway of the thermal control scheme to selectively adjust a temperature of the shot tube plunger. The fluid is discharged through a fluid outlet of the fluid passageway.

Abstract: A preform, for repairing an abradable seal component of a gas turbine engine, includes a multilayer stack that has a first layer and a second layer. The first layer is operative to bond the multilayer stack to a structural substrate of an abradable seal. The first layer includes structural material that corresponds to a material of the structural substrate and braze material that is compatible with the structural material. The second layer includes an abradable material. The multilayer stack is operative to bond to the structural substrate of the abradable seal component during a brazing process such that the second layer forms a replacement abradable layer of the abradable seal component.

Abstract: A gas turbine engine system according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly defined about an axis and a core engine positioned radially inward of the nacelle assembly and having a core passage and at least one core auxiliary duct passage. The at least one core auxiliary duct passage includes an inlet for receiving a portion of a core airflow from the core engine and an outlet for discharging the portion of the core airflow. At least one of the inlet and the outlet are selectively translatable to divert the portion of the core airflow into the at least one core auxiliary duct passage and a mixer disposed between the nacelle assembly and the core engine.

Abstract: A method and apparatus for microplasma spray coating a portion of a turbine vane without masking any portions thereof. The apparatus includes a microplasma gun with an anode, cathode, and an arc generator for generating an electric arc between the anode and cathode. An arc gas emitter injects gas through the electric arc. The electric arc is operable for ionizing the gas to create a plasma gas stream. A powder injector injects powdered material into a plasma stream. A localized area of the turbine vane is coated with the powdered material without having to mask the turbine vane.

Abstract: Non-stick fixtures for selectively masking portions of a workpiece during application of a workpiece coating are described herein. These fixtures have predetermined surfaces thereon having an average surface roughness of about 25 Ra or less and a Rockwell hardness of about 65 Rc or more. The controlled average surface roughness ensures that these fixtures are non-stick with respect to the workpiece coating being applied to the workpieces disposed therein. The controlled Rockwell hardness ensures that the desired average surface roughness can be maintained throughout repeated use of the fixture in harsh coating environments. These fixtures reduce the workpiece coating bridging that occurs between the fixture and the workpiece, and also reduce the amount of overspray that occurs on the workpiece, thereby minimizing the amount of handwork and/or rework that is necessary after the workpiece is coated. This improves process cycle times and yields significantly.