Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having an outer surface and an inner surface that axially extend between a leading edge portion and a trailing edge portion. At least one augmentation feature is disposed on at least the leading edge portion or the trailing edge portion of the outer surface of the platform.

Abstract: A fluid injection system for a gas turbine engine may comprise a fluid injector configured to inject a fluid into an exhaust flow exiting a turbine section of the gas turbine engine. The fluid injector may be coupled to a turbine exit guide vane located at a forward end of an exhaust system of the gas turbine engine. The fluid may decrease a temperature of the exhaust flow exiting the turbine section and/or increase a thrust of the gas turbine engine.

Abstract: A gas turbine engine includes a low speed spool mechanically interconnecting a low pressure turbine and at least one of a fan and a prop, a high speed spool mechanically interconnecting a high pressure turbine and a high pressure compressor, and an epicyclic gear system mechanically coupled to the high speed spool. The gas turbine engine also includes a low pressure compressor mechanically coupled to the high speed spool via the epicyclic gear system. The low pressure compressor may be mechanically independent of the low speed spool. The gas turbine engine may include a plurality of motor-generators for transferring power between the high speed spool and the low pressure compressor.

Abstract: A component for a gas turbine engine. The component includes a first multiple of axial standoff ribs that extend from the first sidewall and a second multiple of axial standoff ribs that extend from the second sidewall. The structural rib that extends between the first multiple of axial standoff ribs and the second multiple of axial standoff ribs.

Abstract: A component for a gas turbine engine includes a interleaved structural rib that extends between the first multiple of axial standoff ribs and the second multiple of axial standoff ribs. The second multiple of structural rib span segments that extend from the second sidewall, the first multiple of structural rib span segments interleaved with the second multiple of structural rib span segments to form an interleaved structural rib that extends between the first sidewall and the second sidewall.

Abstract: An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber. A tangential onboard injector (TOBI) may be disposed in the first airflow path. A radial onboard injector (ROBI) may be disposed in the second airflow path.

Abstract: An airfoil is provided. The airfoil may comprise a cross over, an impingement chamber in fluid communication with the cross over, and a first trip strip disposed on a first surface of the impingement chamber. A cooling system is also provided. The cooling system may comprise an impingement chamber, a first trip strip on a first surface of the impingement chamber, and a second trip strip on a second surface of the impingement chamber. An internally cooled engine part is further provided. The internally cooled part may comprise a cross over and an impingement chamber in fluid communication with the cross over. The cross over may be configured to direct air towards a first surface of the impingement chamber. A first trip strip may be disposed on the first surface of the impingement chamber.

Abstract: A gas turbine engine includes an engine case along an engine axis, a conformal accessory drive gearbox housing mounted to the engine case, and at least one accessory mounted to the conformal housing

Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a tip shelf located in a surface of the tip proximate to a pressure side of the airfoil, wherein the tip shelf has a ledge portion extending from the pressure side to a wall portion extending upwardly from the ledge portion to the tip and wherein the wall portion is configured to have a convex portion with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.

Abstract: A process for directional solidification of a cast part comprises energizing a primary inductive coil coupled to a chamber having a mold containing a material; generating an electromagnetic field with the primary inductive coil within the chamber, wherein said electromagnetic field is partially attenuated by a susceptor coupled to said chamber between said primary inductive coil and said mold; determining a magnetic flux profile of the electromagnetic field after it passes through the susceptor; sensing a component of the magnetic flux in the interior of the susceptor proximate the mold; positioning a mobile secondary compensation coil within the chamber; generating a control field from a secondary compensation coil, wherein said control field controls said magnetic flux; and casting the material within the mold.

Abstract: A thermal management system for a gas turbine engine includes an additively manufactured outer engine duct structure, at least a portion of an internal surface of the additively manufactured outer engine duct structure forming an additively manufactured heat exchanger

Abstract: An aero damping measurement system is provided. The system includes a shroud defining a tunnel, a hub disposed within the tunnel, and a plurality of blades coupled to the hub. The blades may rotate about the hub. A gas pressure probe may have a tip extending to the tunnel to deliver a pressurized burst into the tunnel. An aeromechanical identification system may include a pressurized gas source, a valve in fluid communication with the pressurized gas source, and the gas pressure probe may be in fluid communication with the valve. The valve may control a flow of a pressurized gas from the pressurized gas source into the gas pressure probe. A pressure sensor may be coupled to the gas pressure probe and configured to measure a pressure within the gas pressure probe.

Abstract: A thermal management system for a gas turbine engine includes an additively manufactured nacelle component, at least a portion of the additively manufactured nacelle component forming an additively manufactured heat exchanger that extends into a fan bypass flow.

Abstract: A gas turbine engine may include a rotor overspeed protection (ROP) assembly. The ROP assembly may include an annular blade outer air seal (BOAS) assembly including a ROP segment. The ROP assembly may include a stator vane coupled with the BOAS assembly/. The stator vane may include a stator flange disposed about a forward edge portion of the stator vane. The ROP segment may include a ROP flange extending in an axially aft direction from a main body of the ROP segment toward the stator vane, wherein the ROP flange is disposed radially inward of the stator flange.

Abstract: A gas turbine engine includes a coupling and an air seal attachment. The annular coupling extends along a centerline and has coupling teeth that extend axially rearward. The annular air seal attachment includes an air seal at an axially rearward end and air seal attachment teeth at an axially forward end that extend axially forward to interlock with the coupling teeth such that the air seal attachment teeth and the coupling teeth alternate in a circumferential direction.

Abstract: A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face.

Abstract: An embodiment of a method includes retaining a first workpiece and a second workpiece selectively on a workpiece fixture disposed within a deposition chamber. The workpiece fixture includes tooling including a first workpiece holder, a second workpiece holder, and a first hollow wall. The first workpiece is separated from the second workpiece using the first hollow wall. Energy is selectively applied and directed within the deposition chamber, from an energy source toward a first crucible, the first crucible including a plurality of walls defining an upper recess contiguous with, and disposed directly above a first lower recess, at least the upper recess open to an interior of the deposition chamber. During the step of selectively applying and directing energy, a gas valve is controlled to maintain a partial vacuum in the deposition chamber of greater than 2 Pa to control a size and overlap of at least one coating zone formed around each of the at least one workpiece.

Abstract: A variable vane assembly according to an example of the present disclosure includes an actuator. A connection linkage is pivotable by the actuator. A component receives direct input from the connection linkage providing a cam surface. A driving linkage is interfaced with the cam surface. A variable vane is coupled to the driving linkage. The pivoting of the component and cam surface causes the driving linkage to pivot the vane.

Abstract: A gas turbine engine component includes an airfoil defined by a leading edge, a trailing edge, a pressure sidewall, and a suction sidewall. An internal cavity extends radially through the airfoil and is partially defined by an inner surface of the pressure sidewall and an inner surface of the suction sidewall. A baffle is disposed within the internal cavity, and includes a baffle wall conformal with adjacent surfaces of the internal cavity, a divider separating a forward chamber and an aft chamber, and a plurality of orifices extending through the baffle wall at the forward chamber. At least one axial rib is disposed between the baffle wall and the internal cavity, and is positioned at a discrete spanwise location of the airfoil.

Abstract: A method for joining engine components includes positioning a first plurality of thermal protection structures across a thermal protection space between a first thermal protection surface and a second thermal protection surface. The first and second engine components are locally joined by forming a first plurality of transient liquid phase (TLP) or partial transient liquid phase (PTLP) bonds along corresponding ones of the first plurality of thermal protection structures between the first thermal protection surface and the second thermal protection surface. The second thermal protection surface is formed from a second surface material different from a first surface material of the first thermal protection surface.