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US 11,111,858 B2Granted
Cool core gas turbine engine
AssigneeGeneral Electric Company
InventorsBrandon Wayne Miller, Randy M. Vondrell, Jeffrey Donald Clements, Kurt David Murrow
Life of this patent
Estimating the term…
Abstract
A gas turbine engine includes a compressor section defining a compressor exit temperature, T3. The gas turbine engine also includes a combustion section and a turbine section, with the turbine section defining a turbine inlet temperature, T4. A ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, during operation of the gas turbine engine at a rated speed is less than or equal to 1.85.
Claims18
A gas turbine engine comprising:
- a compressor section defining a compressor exit temperature, T3;
- a combustion section located downstream of the compressor section;
- and
- a turbine section located downstream of the combustion section and defining a turbine inlet temperature, T4, wherein a ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, during operation of the gas turbine engine at a rated speed is less than or equal to 1.85;
- wherein the compressor section, combustion section, and turbine section together define a core air flowpath, wherein the turbine section comprises a first turbine located immediately downstream from the combustion section, wherein the first turbine comprises a plurality of first stage turbine rotor blades, wherein each of the first stage turbine rotor blades extend from a root to a tip and are formed of a wall, wherein the wall of each first stage turbine rotor blade is exposed to the core air flowpath within the turbine section and is configured as a continuous, non-permeable wall to prevent an airflow therethrough such that the first stage turbine rotor blade is an uncooled component.
2 more elements
A method of operating a gas turbine engine comprising a compressor section, a combustion section, and a turbine section in series flow order, the method comprising:
- operating the gas turbine engine at a rated speed such that the turbine section defines a turbine inlet temperature, T4, the compressor section defines a compressor exit temperature, T3, and the gas turbine engine defines a ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, less than or equal to 1.85;
- wherein the compressor section, combustion section, and turbine section together define a core air flowpath, wherein the turbine section comprises a first turbine located immediately downstream from the combustion section, wherein the first turbine comprises a plurality of first stage turbine rotor blades, wherein each of the first stage turbine rotor blades extend from a root to a tip and are formed of a wall, wherein the wall of each first stage turbine rotor blade is exposed to the core air flowpath within the turbine section and is configured as a continuous, non-permeable wall to prevent an airflow therethrough such that the first stage turbine rotor blade is an uncooled component.
A gas turbine engine comprising:
- a compressor section defining a compressor exit temperature, T3;
- a combustion section located downstream of the compressor section;
- and
- a turbine section located downstream of the combustion section and defining a turbine inlet temperature, T4, wherein a ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, during operation of the gas turbine engine at a rated speed is less than or equal to 1.85;
- wherein the compressor section, combustion section, and turbine section together define a core air flowpath, wherein the turbine section comprises a first turbine located immediately downstream from the combustion section, wherein the first turbine comprises a plurality of first stage turbine nozzles, wherein each of the first stage turbine nozzles extend from a root to a tip and are formed of a wall, wherein the wall of each first stage turbine nozzle is exposed to the core air flowpath within the turbine section and is configured as a continuous, non-permeable wall to prevent an airflow therethrough such that the first stage turbine nozzle is an uncooled component.
2 more elements
Dependencies are inferred from the claim text.