Cooling Method for a High-Temperature Radial Gas Turbine Engine
AssigneeTurboGen, LLC
InventorDavid LIOR
Life of this patent
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Abstract
A method for cooling a high-temperature radial gas turbine engine increases turbine thermal efficiency and/or extends turbine operational lifetime. A bleed flow path enables cooling air to flow from a compressor outlet and along surfaces of the gas turbine rotors. The amount of cooling increases in proportion to a bleed fraction, which is defined as the ratio of mass flow in the bleed flow path to total mass flow in the compressor outlet. The heated air in the bleed flow path is mixed with the main mass flow into the turbine engine, so as to restore mass flow into the turbine, while maintaining a high turbine operating temperature and thermal efficiency. The thermal efficiency of a recuperator also increases in proportion to the bleed fraction.
Claims9
A method for cooling hot components of a high-temperature radial gas turbine engine comprising:
- providing a radial gas turbine engine comprising a compressor outlet and at least one rotor comprising a rotor disc and a rotor blade;
- and
- providing a bleed flow path in which a fluid passes from the compressor outlet and flows along surfaces of the rotor disc and rotor blade;
- wherein a ratio of a mass flow in the bleed flow path to a mass flow in the compressor outlet is equal to a bleed fraction which is greater than or equal to zero and less than or equal to one.
Dependencies are inferred from the claim text.