Patent ID: 7004720

Claim:
A high pressure turbine vane assembly for a gas turbine engine, the vane assembly comprising a plurality of airfoils radially extending between inner and outer platforms defining an annular gas path therebetween, at least one of the inner and outer platforms being sealingly engaged to a downstream end of a combustion chamber wall enclosing a combustion chamber of the gas turbine engine in fluid flow communication with said annular gas path, said at least one of said inner and outer platforms defines an inner surface facing said annular gas path and an outer surface which encloses at least part of a cooling air cavity defined between the combustion chamber wall and said at least one of said inner and outer platforms, said outer surface being directly exposed to cooling airflow within said cooling air cavity. wherein a plurality of holes extend between said outer surface and said inner surface within a region of said at least one of the inner and outer platforms substantially intermediate adjacent airfoils, the holes providing fluid flow communication between the cooling air cavity and the annular gas path and directing cooling airflow therethrough such that effusion cooling of the vane assembly is provided.