Patent ID: 7708525
Filing Date: 2010-05-04
Classification: F01D,F05D

Abstract:
1. A gas turbine blade assembly comprising: a neck defining a neck cavity, and having a first end and a second end at an opposite side relative to the first end; a platform having first and second sides, the first side being disposed on and facing the second end of the neck; and an airfoil supported on the second side of the platform, wherein the neck, the platform and the airfoil cooperate to define at least one inner cooling passage extending from the first end to the second end of the neck and through the platform and into the airfoil, the neck defining at least one core channel extending between the at least one cooling passage and the neck cavity, wherein a tube is brazed into the core channel and used to direct flow and effect impingement upon the underside of the platform, and the platform defining at least one film cooling channel extending from a portion of the first side facing the neck cavity to a portion of the second side disposed exterior to the airfoil to permit cooling air to flow through the at least one inner cooling passage into the neck cavity and through a portion of the platform exterior to the airfoil, wherein the at least one film cooling channel has a central longitudinal axis oriented to direct a flow of air toward a trailing edge of the airfoil at an angle of about 30 degrees or about 43 degrees relative to a forward edge of the platform, the forward edge on a same side as a concave side of the airfoil and extending between the trailing edge and a leading edge of the platform.