Provided is a gas turbine vane and blade assembly in which lattice structures are installed between an impingement plate and an effusion plate. The gas turbine vane and blade assembly is capable of enhancing cooling efficiency in an impingement/effusion cooling technique.

In addition, the gas turbine vane and blade assembly can be manufactured using an additive manufacturing technique, and the lattice structures are capable of replacing supports that are used during an additive manufacturing process, and improving not only structural rigidity and stability but also cooling performance.

A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner diameter to an outer diameter and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a hollow airfoil including a plurality of cavities positioned in the airfoil, an insert positioned in one or more of the plurality of cavities of the hollow airfoil, a plurality of cover plates, at least one of which is positioned at one of the inner diameter and at the outer diameter, and a plurality of impingement pans, at least one of which is positioned at one of the inner diameter and at the outer diameter.

A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.

A turbojet engine nacelle having a cascade-type thrust reverser includes a movable frame surrounding a cold air flow path, guided axially by longitudinal rails fixed to an intermediate casing surrounding the cold air flow path and supporting cascades of thrust-reversal guide vanes and mobile rear cowls. The nacelle has a passage of ancillaries opening radially towards the outside of the movable frame, in a zone with no bearing structure passing radially across the cold air flow path. The movable frame is cut in an axial plane at the passage for the passage of ancillaries, and each edge of the cut of the movable frame is connected to the intermediate casing by a slide rail fitted on a longitudinal rail arranged each on one side of this passage.

A gas turbine engine guide vane heat exchanger has guide vane heat exchanger including electroformed fluid channels in electroformed heat exchanger tubes or a heat exchanger core disposed within airfoil. Non-flammable heat conducting liquid or non-metallic foam may fill space between tubes or core and airfoil. Fluid circuit may include channels within electroformed heat exchanger tubes or the heat exchanger core and extend from inlet manifold to outlet manifold for directing fluid or oil through channels and include fluid or oil supply inlet connected to inlet manifold for receiving the fluid or oil flowed into inlet manifold and a fluid or oil supply outlet connected to fluid or oil supply outlet for discharging fluid or oil flowed out of fluid or oil outlet manifold. Heat exchanger tubes or heat exchanger core, inlet manifold, outlet manifold, supply inlet and supply outlet may be integrally and monolithically electroformed together.

An apparatus and method regarding an airfoil for a turbine engine, the airfoil comprising an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further comprising at least one cooling passage extending radially within the interior and defining a primary cooling airflow; and at least one cooling hole having an inlet defining a first cross-sectional area, the inlet in communication with the cooling passage, an outlet defining a second cross-sectional area greater than the first cross-sectional area; wherein the primary cooling airflow enters the inlet in a first direction and exits the outlet in a second direction different than the first direction.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate. A meter section of a cooling aperture is formed in the substrate. An internal coating is applied onto a surface of the meter section. An external coating is applied over the substrate. A diffuser section of the cooling aperture is formed in the external coating and the substrate to provide the cooling aperture.

The present application provides an exhaust frame differential cooling system of a gas turbine engine to mitigate a temperature differential along a compressor and/or a turbine to minimize centerline eccentricity of a shaft. The exhaust frame differential cooling system may include a number of compressor temperature sensors positioned about the compressor and/or a number of turbine temperature sensors positioned about the turbine, an exhaust frame including an inner barrel with a bearing tunnel for the shaft, an outer barrel, and a number of struts extending from the inner barrel to the outer barrel, a blower, and a cooling air metering system that provides cooling air from the blower to the bearing tunnel and through the inner barrel, the struts, and the outer barrel in response to the temperature differential being determined along the compressor and/or the turbine.

A mid-turbine frame (MTF) assembly having: an outer case circumferentially extending around a central axis; an outer ring secured to the outer case and disposed radially inwardly of the outer case relative to the central axis; an inner case structurally connected to the outer case and disposed radially inwardly of the outer ring relative to the central axis; a main plenum circumferentially extending around the central axis and located between the outer case and the outer ring, the main plenum having an inlet fluidly connectable to a source of cooling air, a first outlet fluidly connected to a secondary plenum between the main plenum and the inner case, a second outlet configured to be fluidly connected to a rotor cavity of the low-pressure turbine, and a third outlet configured to be fluidly connected to a plenum surrounding a containment ring of the low-pressure turbine.

A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly includes fan blades defining a corresponding fan area (A.sub.FAN). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and includes radially-extending vanes arranged in a circumferential array with at least one vane including a heat transfer element for heat transfer from a first fluid contained within each element to an airflow passing over a surface of each heat transfer element before entering the fan assembly inlet. Each heat transfer element extends axially along the corresponding vane, with a swept heat transfer element area (A.sub.HTE) being the wetted surface area of all heat transfer elements in contact with the airflow. A Fan to Element Area parameter F.sub.EA of A.sub.HTE/A.sub.FAN lies in the range of 47 to 132.