Turn caps for airfoils of gas turbine engines having a first pressure-side turn passage extending from a respective inlet to a respective outlet within the turn cap, a first suction-side turn passage extending from a respective inlet to a respective outlet within the turn cap, and a merging chamber fluidly connected to the outlets of the first pressure-side turn passage and the first suction-side turn passage, wherein each of the first suction-side turn passage and the first pressure-side turn passage turn a direction of fluid flow from a first direction to a second direction such that a fluid flow exiting the first suction-side turn passage and the first pressure-side turn passage are aligned when entering the merging chamber.

Turn caps for airfoils of gas turbine engines having a first pressure-side turn passage extending from a respective inlet to a respective outlet within the turn cap, a first suction-side turn passage extending from a respective inlet to a respective outlet within the turn cap, and a merging chamber fluidly connected to the outlets of the first pressure-side turn passage and the first suction-side turn passage, wherein each of the first suction-side turn passage and the first pressure-side turn passage turn a direction of fluid flow from a first direction to a second direction such that a fluid flow exiting the first suction-side turn passage and the first pressure-side turn passage are aligned when entering the merging chamber.

An airfoil (10) for a turbine engine includes an array of features (22) positioned in an interior portion (11) of the airfoil (10). Each feature (22) extends from a pressure (14) side to a suction side (16). The array includes multiple radial rows (A-N) of features (22) with the features (22) in each row (A-N) being interspaced radially to define coolant passages (24) therebetween. The radial rows (A-N) are spaced along a forward-to-aft direction toward an airfoil trailing edge (20). The coolant passages (24) of the array are fluidically interconnected to lead a pressurized coolant toward the trailing edge (20) via a serial impingement on to the rows of features (22). The coolant passages (24) are geometrically configured to bias a coolant flow therethrough toward a first side (14) in relation to a second side (16) of the outer wall (12) to effect a greater cooling of the first side (14) than the second side (16).

A method and apparatus for a turbine engine having a compressor section, combustion section, and a turbine section in an axial flow arrangement with a cooling circuit in fluid communication with at least one of the compressor section, combustion section, or turbine section. The method and apparatus further including separating particles from a cooling air that flows through the cooling circuit.

An apparatus and method for assembling an inducer assembly for inducing a rotation on an airflow passing within a turbine engine. The inducer assembly can provide a volume of fluid from a compressor section to a turbine section of the engine. The inducer assembly can include the combination of separate segments to form an annular inducer.

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 the mixer surrounding the hub, and a plurality of deswirling struts circumferentially spaced apart with respect to a central axis of the TEC and located entirely within an axial length of the mixer. The mixer defines a trailing edge having one or more upstream-most locations thereof where the mixing of the exhausted gases and the bypass air stream begins to take place. The deswirling struts each extend radially across the annular exhaust gas duct and interconnect the mixer and the hub, defining a trailing edge positioned upstream of and axially spaced away from the one or more upstream-most locations of the trailing edge of the mixer.

In one embodiment, a gas turbine engine for mounting to a rear of an aircraft fuselage has a propulsor that rotates on a first axis, and an engine core including a compressor section, a combustor section, and a turbine section, with the turbine section being closer to the propulsor than the compressor section. The engine core is aerodynamically connected to the propulsor and has a second axis. A nacelle is positioned around the propulsor and engine core. The nacelle is attached to the wing of the aircraft. A downstream end of the nacelle has at least one pivoting door with an actuation mechanism to pivot the door between a stowed position and a vertical deployed position in which the door inhibits a flow to provide a thrust reverse of the flow.

In accordance with one aspect of the disclosure, a stream diverter for a gas turbine engine is disclosed. The stream diverter may include a first air duct, a second air duct, a third air duct, and a door operatively associated with the second and third air ducts of the gas turbine engine. The door may have at least an open position allowing air from the second air duct to flow into the third air duct and a closed position preventing air from flowing between the ducts.