An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

A device for cooling a component to be cooled of a gas turbine/turbo machine having a hot-gas-impinged outer surface, a target surface of the component, and an integrated cooling passage, includes: an impingement cooling element arranged within the cooling passage, the impingement cooling element having plural impingement cooling bores; and a surface structure arranged on the target surface. The impingement cooling element is spaced apart from the target surface of the component and configured so as to conduct a cooling fluid as an impingement cooling jet is onto the target surface, such that the impingement cooling jet impinges on the surface structure.

A compressor case to blade tip clearance system comprising a rotor having blades with tips, the case including an inner case comprising at least one surface feature fluidly coupled to a distribution manifold disposed in a cooling air passageway, the at least one surface feature configured to interact with the cooling air, and a tip clearance located between the tips and the inner case; wherein the tip clearance is maintained responsive to a flow of the cooling air over the at least one surface feature.

A turbine vane, a turbine, and a gas turbine capable of reducing thermal stress are provided. The turbine vane may include an airfoil including a leading edge and a trailing edge, an inner shroud disposed at one end of the airfoil to support the airfoil, an outer shroud disposed at the other end of the airfoil to support the airfoil and configured to face the inner shroud, a first cooling passage and a second cooling passage configured to extend in a height direction thereof, and a first passage bending part configured to connect the first cooling passage and the second cooling passage, and the first passage bending part is positioned inside the inner shroud or the outer shroud.

An air-to-air heat exchanger in flow communication with a gas turbine engine is provided. The air-to-air heat exchanger has an air-to-air heat exchanger potential defined by a product raised to a half power, the product being a heat transfer surface area density associated with the air-to-air heat exchanger multiplied by an airflow conductance factor associated with the gas turbine engine. The air-to-air heat exchanger potential is between about 6.7 and 19.5 for a bypass ratio associated with the gas turbine engine between about 3 and 10 and the heat transfer surface area density being between about 3,000 m.sup.2/m.sup.3 and 10,000 m.sup.2/m.sup.3 and is between about 2.9 and 12.2 for a bypass ratio associated with the gas turbine engine between about 10 and 20 and the heat transfer surface area density being between about 1,000 m.sup.2/m.sup.3 and 10,000 m.sup.2/m.sup.3.

A high-temperature component according to an embodiment is a high-temperature component which requires cooling by a cooling medium, and includes: a plurality of cooling passages through which the cooling medium is able to flow; a header portion to which downstream ends of the plurality of first cooling passages are connected; and at least one outlet passage for discharging the cooling medium flowing into the header portion to outside of the header portion. A roughness of an inner wall surface of the at least one outlet passage is not greater than a roughness of an inner wall surface of the plurality of first cooling passages in a region where a flow-passage cross-sectional area of the outlet passage is the smallest.

A turbine vane assembly adapted for use with a gas turbine engine includes an airfoil and a spar. The airfoil is formed to define a cavity that extends into the airfoil. The spar is located in the cavity to define a cooling passage that extends around the spar between the spar and the airfoil. The turbine vane assembly includes cooling features to aid heat transfer of the turbine vane assembly during operation in the gas turbine engine.

The present invention provides a heat transferring device and a method for making thereof. The heat transferring device has a thermal conducting substrate and a porous layer. The thermal conducting substrate has a plurality of protrusions and concave bottom surfaces. The concave bottom surfaces are located between the protrusions. The porous layer is embedded between the protrusions. The present invention also provides a high temperature material transferring system comprising a cylindrical container and the heat transferring device disposed on the surface of the cylindrical container.

A cooling assembly comprises a cooling cavity disposed inside of a turbine assembly. The cooling cavity is configured to direct cooling air inside a body of the turbine assembly. The cooling assembly comprises a cross-bank fluidly coupled with the cooling cavity and positioned to direct at least some of the cooling air out of the cooling cavity and outside of the body. The cross-bank comprises plural pins having first ends coupled with a first side interior surface of the body and opposite second ends coupled with a second side interior surface of the body. The cross-bank also includes a cross-bar connecting the pins. The cross-bar extends between the pins such that the cross-bar has a first end coupled with an exterior surface of a first pin of the pins and an opposite second end coupled with an exterior surface of a second pin of the pins.

A component for a gas turbine engine includes a mateface with a purge flow interface, the mateface comprises a pocket located in communication with a feather seal slot in a mateface. A vane for a gas turbine engine includes a platform that extends from the airfoil, the platform comprising a mateface with a feather seal slot and a pocket in communication with the feather seal slot, wherein the pocket is of a cross-sectional shape larger than the feather seal slot.