Cooling system for a turbine engine
AssigneeGeneral Electric Company
InventorsBrandon Wayne Miller, Jeffrey Anthony Hamel
Life of this patent
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Abstract
A gas turbine engine including a compressor section and a turbine section, coupled by one or more shafts, is provided. The compressor section progressively compresses air and includes an aft-stage of rotor blades rotatable about an axial direction of the gas turbine engine. A cooling system is included with the gas turbine engine for cooling compressed air in or from the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and one or more fluid lines in fluid communication with the fluid tank. The one or more fluid lines include an outlet positioned adjacent to the aft-stage of rotor blades for injecting cooling fluid into the compressed air proximate the aft-stage of rotor blades.
Claims17
A gas turbine engine defining an axial direction, the gas turbine engine comprising:
- a compressor section for progressively compressing air, the compressor section including an aft-stage of rotor blades rotatable about the axial direction;
- a cooling system for cooling compressed air in or from the compressor section, the cooling system comprising
- a fluid tank for storing a volume of cooling fluid;
- and
- one or more fluid lines in fluid communication with the fluid tank, the one or more fluid lines including an outlet positioned adjacent to the aft-stage of rotor blades of the compressor section for injecting cooling fluid in an upstream direction into the compressed air proximate the aft-stage of rotor blades;
- and
- wherein the gas turbine engine further defines a radial direction, wherein the compressor section further includes a diffuser positioned aft of the aft-stage of rotor blades, and wherein the one or more fluid lines include a fluid line attached to a stationary component inward of the diffuser along the radial direction.
A cooling system for cooling compressed air in or from a compressor section of an aeronautical gas turbine engine, the cooling system comprising:
- a fluid tank for storing a volume of cooling liquid;
- and
- one or more fluid lines in fluid communication with the fluid tank, the one or more fluid lines including an outlet and configured to be mounted such that the outlet is positioned adjacent to an aft-stage of rotor blades of the compressor section for injecting cooling liquid in an upstream direction into compressed air proximate the aft-stage of rotor blades of the compressor section;
- and
- wherein the aeronautical gas turbine engine further defines a radial direction, wherein the compressor section further includes a diffuser positioned aft of the aft-stage of rotor blades, and wherein the one or more fluid lines include a fluid line attached to a stationary component inward of the diffuser along the radial direction.
Dependencies are inferred from the claim text.