Cool core gas turbine engine
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.
1. 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. The gas turbine engine of claim 1 , wherein the compressor exit temperature, T3, is greater than about 1,200 degrees Rankine and less than about 2,000 degrees Rankine.
3. The gas turbine engine of claim 1 , wherein the ratio, T4:T3, of the turbine inlet temperature, T4, to compressor exit temperature, T3, during operation of the gas turbine engine at the rated speed is greater than or equal to 1.5 and less than or equal to 1.8.
4. The gas turbine engine of claim 1 , wherein the compressor section further defines an overall pressure ratio greater than or equal to 25 during operation of the gas turbine engine at the rated speed.
5. The gas turbine engine of claim 1 , wherein the turbine inlet temperature, T4, in degrees Rankine is at least about five percent (5%) less than a value determined by the formula: 1.88×T3+101 degrees Rankine, wherein T3 is the compressor exit temperature, T3, having a value greater than 1,200 degrees Rankine.
6. The gas turbine engine of claim 1 , wherein the walls forming the first stage turbine rotor blades are each formed of a refractory material, wherein the gas turbine engine is configured to generate at least about 20,000 pounds of thrust during operation at the rated speed, and wherein the gas turbine engine is configured as an aeronautical turbofan engine.
7. 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.
8. The method of claim 7 , wherein the compressor exit temperature, T3, is greater than about 1,200 degrees Rankine and less than about 2,000 degrees Rankine.
9. The method of claim 7 , wherein operating the gas turbine engine at the rated speed comprises operating the compressor section of the gas turbine engine to define an overall pressure ratio greater than or equal to 25.
10. The method of claim 7 , wherein operating the gas turbine engine at the rated speed comprises operating the gas turbine engine to define a turbine inlet temperature, T4, in degrees Rankine, less than about five percent (5%) of a value determined by the formula: 1.88×T3+101 degrees Rankine, wherein T3 is the compressor exit temperature, T3, having a value greater than 1,200 degrees Rankine.
11. The method of claim 7 , wherein operating the gas turbine engine at the rated speed comprises operating the gas turbine engine to define a turbine inlet temperature, T4, in degrees Rankine, about ten percent (10%) less than a value determined by the formula: 1.88×T3+101 degrees Rankine, wherein T3 is the compressor exit temperature, T3, having a value greater than 1,200 degrees Rankine.
12. The method of claim 7 , wherein the walls forming the first stage turbine rotor blades are each formed of a refractory material, wherein the gas turbine engine is configured to generate at least about 20,000 pounds of thrust during operation at the rated speed, and wherein the gas turbine engine is configured as an aeronautical turbofan engine.
13. 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.
14. The gas turbine engine of claim 13 , wherein the compressor exit temperature, T3, is greater than about 1,200 degrees Rankine and less than about 2,000 degrees Rankine.
15. The gas turbine engine of claim 13 , wherein the compressor section further defines an overall pressure ratio greater than or equal to 25 during operation of the gas turbine engine at the rated speed.
16. The gas turbine engine of claim 13 , wherein the turbine inlet temperature, T4, in degrees Rankine is at least about five percent (5%) less than a value determined by the formula: 1.88×T3+101 degrees Rankine, wherein T3 is the compressor exit temperature, T3, having a value greater than 1,200 degrees Rankine.
17. The gas turbine engine of claim 13 , wherein each of the first stage turbine nozzles is configured without any cooling holes defined there.
18. The gas turbine engine of claim 13 , wherein the walls forming the first stage turbine nozzles are each formed of a refractory material, wherein the gas turbine engine is configured to generate at least about 20,000 pounds of thrust during operation at the rated speed, and wherein the gas turbine engine is configured as an aeronautical turbofan engine.