Abstract: A clearance control system for a gas turbine engine comprises at least one case support associated with an engine case defining an engine center axis. A clearance control ring is positioned adjacent the at least one case support to form an internal cavity between the engine case and the clearance control ring. The clearance control ring includes a first mount feature. An outer air seal has a second mount feature cooperating with the first mount feature such that the clearance control ring can move independently of the engine case in response to changes in temperature. An injection source inject flow into the internal cavity to control a temperature of the clearance control ring to allow the outer air seal to move in a desired direction to maintain a desired clearance between the outer air seal and an engine component. A gas turbine engine and a method of controlling tip clearance in a gas turbine engine are also disclosed.

Abstract: A vane stage includes a ringcase extending circumferentially about a center axis of the vane stage. The ringcase extends completely about the center axis to form a first ring. An inner shroud extends circumferentially about the center axis of the vane stage. The inner shroud extends completely about the center axis to form a second ring positioned radially within the ringcase relative the center axis. A plurality of stationary half vanes extend radially between the ringcase and the inner shroud, and are circumferentially spaced about the center axis. The plurality of stationary half vanes are integral with the ringcase and the inner shroud.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan situated at an inlet of a bypass passage, and a core engine configured to drive the fan. The core engine includes a low pressure compressor section driven by a low pressure turbine section, and a high pressure compressor section driven by a high pressure turbine section. The fan has a fan diameter, Dfan, and the high pressure compressor section has a compressor diameter, Dcomp. The fan diameter Dfan and the compressor diameter Dcomp have an interdependence represented by a scalable ratio Dfan/Dcomp that is greater than about 4.5.

Abstract: An air foil includes a body and a filler. The body may include a chamber formed therein and includes a first metallic material. The filler may at least partially occupy the chamber of the body and may include a second metallic material. The second metallic material may have a comparatively more negative electrode potential than the first metallic material.

Abstract: A gas turbine engine includes a fan, a compressor, a combustor, a turbine, a bypass duct, and a bearing compartment assembly. The bearing compartment assembly includes a fluid pump, a compartment, a fluid line between the fluid pump and the compartment, and a damper check valve located in the fluid line. The damper check valve is a unitary, monolithic component that is configured to restrict a reverse flow from the compartment to the fluid pump substantially more than the damper check valve restricts a standard flow from the fluid pump to the compartment.

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 rotatable component and a non-rotatable component. A seal is carried on one of the rotatable component or the non-rotatable component to provide sealing there between. The seal includes a geopolymer seal element.

Abstract: A process includes solution heat treating a nickel based superalloy with greater than about 40% by volume of gamma prime precipitate to dissolve the gamma prime precipitate in the nickel based superalloy; cooling the nickel based superalloy to about 85% of a solution temperature measured on an absolute scale to coarsen the gamma prime precipitate such that a precipitate structure is greater than about 0.7 micron size; and wrought processing the nickel based superalloy at a temperature below a recrystallization temperature of the nickel based superalloy. A material includes a nickel based superalloy with greater than about 40% by volume of gamma prime precipitate in which the precipitate structure is greater than about 0.7 micron size.

Abstract: A cooling system for a gas turbine engine may comprise a heat exchanger configured to receive a cooling airflow. The heat exchanger may receive the cooling flow at a cooling flow input of the heat exchanger and output the cooling airflow at an exhaust output of the heat exchanger. An exhaust distribution manifold may be fluidly coupled to the exhaust output of the heat exchanger. The exhaust distribution manifold may define a plurality of openings.

Abstract: A gas turbine engine includes a stator vane. A blade outer air seal may be adjacent to the stator vane. A seal assembly may be disposed between the blade outer air seal and the stator vane. The seal assembly may include a first seal disposed between a first airflow path and a second airflow path. A second seal may be disposed between the first airflow path and the second airflow path. The second seal may be disposed in a series arrangement with the first seal. The second seal may include a seal support extending radially therefrom which mechanically supports the first seal during engine operation.

Abstract: An egress seal fitting for a bulkhead penetration may comprise a housing, a driver configured to couple to the housing and define a cavity within the housing, and a seal member configured to be disposed within the cavity of the housing, wherein the driver is configured to pass a cable through the housing and apply pressure to the seal member.

Abstract: A method of calibrating a laser measurement device includes positioning a calibration part, having known geometric specifications, at varying distances from the laser measurement device (within a measurement depth-of-field of the laser measurement device). The detected geometric specifications, measured at these different positions, are compared with the known specification(s) of the calibration part in order to improve the accuracy of the calibration. A calibration standoff fixture having a plurality of adjustable legs may be used in conjunction with the method to facilitate changing the position of the calibration part relative to the laser measurement device.

Abstract: A rotary tool includes a shank, cutting portion, coolant inlet, coolant outlet, and channel system. The cutting portion is connected to and extends from the shank. The cutting portion includes a cutting edge. The coolant inlet and the coolant outlet are disposed in the shank. The channel system is contained in the rotary tool and is a closed circulation loop system such that the coolant is contained within the channel system as the coolant is circulated within the rotary tool. The channel system includes a delivery path and a return path. The delivery path is fluidly connected to the coolant inlet and includes a shape corresponding to a shape of the cutting edge. The return path is fluidly connected to the coolant outlet and to the delivery path at a location in the cutting portion of the rotary tool.

Abstract: An airfoil for a gas turbine engine includes axial flow and radial flow cooling circuits defined within an airfoil body. A baffle disposed in spaced relation to an inner surface of the airfoil has a plurality of impingement cooling holes configured to direct a cooling fluid at an inner surface of the airfoil body and an axial extent from the leading edge defined by an aft wall, with the axial extent being substantially constant between the inner and outer end walls and defined by a plane perpendicular to an engine axis. A first radially-extending rib is angled with respect to the baffle to define a first passage between the first rib and the baffle that tapers in cross-sectional area between the inner end wall and the outer end wall, becoming larger in cross-sectional area in a direction of cooling fluid flow through the first passage.

Abstract: An embodiment of a tool assembly includes a spring-loaded shaft assembly disposed along a first axis, a hub, and a roller disk. The hub is connected to a distal end of the spring-loaded shaft assembly. The hub has an upper hub portion adjacent to the distal end of the spring-loaded shaft assembly aligned with the first axis, and a lower hub portion extending along a second axis. The second axis forms an angle relative to the first axis. The roller disk is joined to the lower portion of the hub, and has a working surface about its perimeter. The roller disk is rotatable about the second axis parallel to the second portion of the hub. The working surface includes a profile along its width such that an effective radius of the roller disk varies along a width thereof.

Abstract: An airfoil includes a core having a first surface, a skin having a second surface disposed over at least a portion of the first surface of the core, and at least one of a transient liquid phase (TLP) bond and a partial transient liquid phase (PTLP) bond. The bond(s) are disposed between the first surface and the second surface, joining the skin to the core.

Abstract: A sensor system for use in a gas turbine engine. The sensor system includes a reflective surface configured to reflect optical features corresponding to combustion in a combustion chamber that exit the combustion chamber via at least one opening therein. The sensor system further includes an optical sensor configured to receive the reflected optical features from the reflective surface.

Abstract: A turbine engine component may comprise a Ceramic Matrix Composite (CMC) structure including a plurality of nominally dense plies, wherein each of the plurality of the nominally dense plies are bonded by at least one of a Field Assisted Sintering Technique (FAST), a Spark Plasma Sintering (SPS), or a localized heating at a bonding interface. The turbine engine component may include an airfoil extending between a first platform and a second platform, wherein the airfoil, the first platform, and the second platform define the CMC structure.

Abstract: A cooling scheme for a blade outer air seal includes a perimeter cooling arrangement configured to convectively cool a perimeter of the blade outer air seal, and a core cooling arrangement configured to cool a central portion of the blade outer air seal through impingement cooling and to provide film cooling to an inner diameter face of the blade outer air seal.

Abstract: A seal plate is disclosed herein that is annular in shape and configured to rotate about a centerline. The seal plate includes a seal body having a first axial end forming a sealing surface and a second axial end opposite the first axial end, a slinger ring axially rearward of the seal body and having a plurality of radially extending tabs separated by a plurality of slots, and a groove between the seal body and the slinger ring. The tabs on the slinger ring are configured to direct lubricant at least partially radially inward into the groove and at least partially in a direction of rotation of the seal plate to cool the seal body.