Abstract:
PROBLEM TO BE SOLVED: To provide an airfoil having an improved cooling mechanism. SOLUTION: An airfoil part of a turbine includes internal cavities 52, 54 and a plurality of indentations 90 disposed on inside surfaces 38 of the internal cavities. By the dents 90, heat transfer is improved to cool the internal cavities 52, 54 of the airfoil to elongate the service life of the airfoil, and the amount of compressor bleed air required is reduced to optimize engine efficiency. The advantage of such a cooling mechanism is that it does not restrict a cooling flow in the internal cavities 52, 54. The indentations 90 may have variable patterns or different geometric shapes.
Abstract:
A heat shield is disclosed. The heat shield may comprise a body having a back surface and an opposite front surface, wherein an opening in the body communicates through the front and back surfaces. The heat shield may further comprise at least one radial rail disposed on the back surface and extending radially outward from the opening for directing cooling air flow.
Abstract:
An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall, which includes a shell, a heat shield and an annular body. The body at least partially defines a first aperture through the shell and the heat shield. The body also defines one or more second apertures through which air is directed into the first aperture and provides non-uniform cooling to the body.
Abstract:
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.
Abstract:
A liner panel for a combustor of a gas turbine engine includes a multiple of heat transfer augmentors. At least one of the multiple of heat transfer augmentors includes a hemi-spherical protuberance.
Abstract:
A liner panel is provided for a combustor of a gas turbine engine. The liner panel includes a multiple of heat transfer augmentors. At least one of the multiple of heat transfer augmentors includes a depression, where the depression includes an entrance to at least one passage through the liner panel.
Abstract:
A liner panel for use in a combustor of a gas turbine engine includes a nozzle includes an inner periphery along an axis. The inner periphery includes a flow guide around the axis. A wall assembly for use in a combustor of a gas turbine engine includes a support shell with a first inner periphery along an axis. The wall assembly also includes a liner panel with a second inner periphery along the axis, the second inner periphery including a spiral flow guide around the axis. A method of reducing recirculation into a dilution passage in a combustor liner panel of a gas turbine engine includes contouring a dilution passage to match a natural vena contracta of a fluid flowing therethrough.
Abstract:
A shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall.
Abstract:
A liner panel for a combustor of a gas turbine engine includes a multiple of heat transfer augmentors which extend from a cold side thereof. At least one of the multiple of heat transfer augmentors includes a first heat transfer augmentation feature with a second heat transfer augmentation feature stacked thereon.