Patent ID: 7458766
Filing Date: 2008-12-02
Classification: F01D

Abstract:
1. In a gas turbine engine, a turbine blade cooling system comprises a compressor having a stage of compressor blades and compressed air bleed means, a stage of turbine blades downstream of said compressor, a disk, said stage of turbine blades being supported on said disk, a cowl covering the downstream face of said disk in spaced, co-axial relationship therewith, bled compressor air diffusion means connected in flow series with said space, holes in the rim of said disk, which holes connect said space with the roots of said turbine blades, and wherein the inner surface of said cowl is formed so as to pressurize said diffused air to a magnitude appropriate to the cooling flow requirements of said turbine blades wherein the cowl is vaned to define a radial compressor.