Turbine blade cooling system
AssigneeROLLS-ROYCE plc
InventorsGeoffrey M Dailey, Guy D Snowsill
Life of this patent
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Abstract
Efficient cooling of a stage of gas turbine engine turbine blades (36) is achieved by first reducing the pressure of the cooling air after it has been bled from the annulus of the compressor (12) by passing it through a diffuser (30), to a pressure magnitude lower than is required at entry to the turbine blades, then re-pressurizing the bled air up to the required entry pressure, by passing it through a radial compressor defined by a cowl (44) positioned in close spaced, co-rotational relationship with the downstream face of the associated turbine disk (34).
Claims10
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.
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 said bled compressor air diffusion means comprises a radial turbine.
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 said compressor air bleed means comprises holes in the outer wall of the compressor annulus.
A method of cooling a stage of gas turbine engine turbine blades comprising the steps of first reducing the pressure of cooling air bled from an associated compressor by passing it through a diffuser so as to achieve a pressure lower than is required at entry to the turbine blades, then re-pressurizing said bled air up to the required entry pressure by passing it through a radial compressor defined by a vaned cowl positioned in close spaced, co-rotational rotational relationship with the downstream face of the associated turbine disk.
Dependencies are inferred from the claim text.