Patent ID: 7607891
Filing Date: 2009-10-27
Classification: F01D,F05D,Y02T

Abstract:
1. A component for a gas turbine engine comprising: an airfoil extending from a radially inner portion to a radially outer portion and having spaced pressure walls with said airfoil being generally hollow between said walls, and said airfoil extending from a leading edge to a trailing edge; and internal cooling flow passages formed in said airfoil, there being a direct cooling passage directing cooling air from said radially inner portion toward said radially outer portion of said airfoil, said direct cooling passage directing cooling air through a plurality of metering openings to skin cooling openings near said trailing edge of said airfoil, and a serpentine cooling path positioned toward said leading edge from said direct cooling passage, said serpentine cooling path directing air radially outwardly toward said radially outer portion of said airfoil, then back radially inwardly, and eventually back radially outwardly, said direct cooling passage being separated from a portion of said serpentine cooling path by a divider wall, said divider wall having a tip flag at a radially outer location which extends toward said trailing edge, and which selectively communicates a portion of air in the serpentine cooling path toward said trailing edge along a tip flag path defined radially outwardly of said tip flag, and a first pedestal array extending between said walls in said tip flag path, and along said tip flag, to direct heat into cooling air flowing along said tip flag path.