A combustor pipe is linked to a vane shroud in which a vane is provided, and includes an inlet, an outlet, an inner pipe of which an inner space is a flow path for passing a combustion gas, a first cooling flow path through which a cooling medium passes being formed inside a wall that forms the flow path; and an outer pipe on an outer circumference of the inner pipe and secured to the inner pipe. A second cooling flow path through which a cooling medium passes and which is connected to the first cooling flow path near the outlet of the combustor pipe is formed between an outer circumferential surface of the inner pipe and an inner circumferential surface of the outer pipe, and a cooling promoting structure is formed on the outer pipe, inside the second cooling flow path near the first cooling flow path.

A hot section part of a turbine engine configured to be exposed to hot gases includes a first wall, a second wall, a plurality of pedestals, a plurality of impingement cooling holes, and a plurality of effusion cooling passages. The walls are spaced apart to form an intervening cavity, and each pedestal extends through the intervening cavity. The impingement cooling holes extend through the second wall to admit a flow of cooling air into the intervening cavity. Each effusion cooling passage is associated with a different one of the plurality of pedestals and is disposed at a predetermined angle relative to its associated principal axis. A portion of the flow of cooling air admitted to the intervening cavity is directed through at least a portion of each of the plurality of pedestals and onto the first wall inner surface.

A structure is provided for a turbine engine. The structure includes a shell with a first surface, and a heat shield with a textured second surface and a textured third surface. The texture of a portion of the second surface is different than the texture of a portion of the third surface. The first surface and the second surface define a first cooling cavity between the shell and the heat shield. The first surface and the third surface define a second cooling cavity between the shell and the heat shield.

A heat shield panel for a combustor of a gas turbine engine including a panel body having a first surface and a second surface. The second surface being configured to be oriented toward a combustor liner of the combustor. The heat shield further includes a plurality of first pin fins projecting from the second surface of the panel body. Each of the plurality of first pin fins has a rounded top opposite the second surface. The heat shield further includes a plurality of second pin fins projecting from the second surface of the panel body. Each of the plurality of second pin fins has a flat top opposite the second surface. The plurality of second pin fins are intermittently spaced amongst the plurality of first pin fins. The plurality of second pin fins are organized in a uniform distribution across the second surface of the heat shield panel.

A wall assembly that may be for a combustor of a gas turbine engine includes a liner having a hot face that defines a combustion chamber, an opposite cold face, and a plurality of effusion holes. A shell of the assembly is spaced outward from the cold face and includes a plurality of impingement holes each having a centerline orientated substantially normal to the cold face. A plurality of cooling member arrays of the liner each include a first plurality of members that may be pins projecting outward from the cold face to conduct heat out of the liner. Each array is spaced between adjacent effusion holes and is symmetrically orientated about the respective centerline.

A helicoidal structure promotes cooling of a liner applied in an annular space of a double-shell structure formed of the liner and a flow sleeve to cool a duct assembly, by increasing the residence time and cooling area of cooling compressed air on the surface of the liner. The helicoidal structure includes a helicoidal rib protruding from a surface of the liner to guide the cooling compressed air at a predetermined angle with respect to an axial direction of the liner, the helicoidal rib having cooling holes and forming a first cooling passage along which main-cooling compressed air flows and a second cooling passage along which auxiliary-cooling compressed air flows. The second cooling passage induces a flow of the auxiliary-cooling compressed air through the cooling holes, and the flow of the auxiliary-cooling compressed air is derived from a flow of the main-cooling compressed air in the first cooling passage.

A combustor panel may comprise a panel wall comprising a proximal surface and a distal surface. A standoff may be formed over the distal surface of the panel wall. An aperture may be formed through the standoff and may extend from a face of the standoff to the proximal surface of the panel wall. The face of the first standoff may be oriented at an angle between 90 degrees and 160 degrees relative to a plane parallel to the distal surface of the panel wall.

A heat shield panel for use in a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side, a cold side and at least one attachment mechanism having a stud and a central axis extending through the stud and a plurality of standoff pins positioned circumferentially around the stud, the standoff pins having a radial extent, a circumferential extent that is greater than the radial extent, a radially outer surface having a radially convex shape and a radially inner surface having a radially concave shape.

There is disclosed wall cooling system 50 having a double-wall geometry. A first wall 55 and a second wall 60 extend over a plan area with the second wall spaced from the first wall by a gap. The first wall 55 has multiple upstanding members 65 spanning the gap and contacting the second wall 60 such that the first and second walls are mechanically and thermally connected. The first wall 55 is shaped so as to provide a two-dimensional array of crests 85 and recesses 90. The crests 85 are spaced from the second wall 60. The first wall 55 has a plurality of through-holes 70 for flow of coolant through the first wall and into the gap. The cooling system 50 is suitable for use in a gas turbine engine 10, for example in the turbine 17, 19.

A gas turbine engine component having a first surface in communication with a core airflow. The gas turbine engine component further includes a second surface, different than the first surface, for cooling the first surface. The gas turbine engine component further includes a heat transfer augmentor extending from the second surface. The gas turbine engine component further includes a heat transfer augmentor effusion hole extending through the gas turbine engine component from a sidewall of the heat transfer augmentor to the first surface.