Geared gas turbine engine architecture for enhanced efficiency
AssigneeUnited Technologies Corporation
InventorKarl L. Hasel
Life of this patent
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Abstract
An example gas turbine engine includes, among other things, a geared architecture rotatably coupled to the fan drive shaft, and a high pressure compressor. The gas turbine engine is configured so that a core temperature at an exit of the high-pressure compressor is approximately in a range of about 1150 to about 1350 degrees Fahrenheit at take-off. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to a primary stream exhaust velocity, is approximately in a range of about 0.75 to about 0.90. A Bypass Ratio of the engine is greater than about 8.0.
Claims30
A gas turbine engine comprising:
- a fan having a drive shaft, the fan being configured to direct air into a bypass duct and into an engine core;
- a geared architecture rotatably coupled to the fan drive shaft;
- and
- a first compressor section, wherein the gas turbine engine is configured to have a core temperature at an exit of the first compressor section is in a range of about 1150 to about 1350 degrees Fahrenheit at take-off,
- wherein the gas turbine engine is configured so that a ratio of a fan stream exhaust velocity divided by a primary stream exhaust velocity is in a range of about 0.75 to about 0.90,
- wherein a ratio of a volume of air passing into the bypass duct divided by a volume of air passing into the engine core is greater than about 8.0.
A gas turbine engine comprising:
- a fan having a drive shaft, the fan being configured to direct air into a bypass duct and into an engine core;
- a geared architecture rotatably coupled to the fan drive shaft, the geared architecture having a speed reduction ratio that is greater than or equal to about 2.3;
- and
- a first compressor section, wherein the gas turbine engine is configured to have a core temperature at an exit of the first compressor section is in a range of about 1150 to about 1350 degrees Fahrenheit at take-off,
- wherein a ratio of a volume of air passing into the bypass duct divided by a volume of air passing into the engine core is greater than about 8.0,
- wherein the gas turbine engine is configured so that a ratio of a fan stream exhaust velocity divided by a primary stream exhaust velocity, is greater than about 0.75.
A gas turbine engine comprising:
- a fan having a drive shaft, the fan being configured to direct air into a bypass duct and into an engine core;
- a geared architecture rotatably coupled to the fan drive shaft, the geared architecture having a speed reduction ratio that is greater than or equal to about 2.3;
- and
- a first compressor section, wherein the gas turbine engine is configured to have a core temperature at an exit of the first compressor section is in a range of about 1150 to about 1350 degrees Fahrenheit at take-off,
- wherein a ratio of a volume of air passing into the bypass duct divided by a volume of air passing into the engine core is greater than about 8.0,
- wherein the gas turbine engine is configured so that a ratio of a fan stream exhaust velocity divided by a primary stream exhaust velocity, is less than about 0.90.
A gas turbine engine comprising:
- a fan having a drive shaft, the fan being configured to direct air into a bypass duct and into an engine core;
- a geared architecture rotatably coupled to the fan drive shaft, the geared architecture having a speed reduction ratio that is approximately in a first range of about 2.3 to about 4.2,
- wherein the gas turbine engine is configured so that a ratio of a fan stream exhaust velocity divided by a primary stream exhaust velocity, is in a second range of about 0.75 to about 0.90 when cruising at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.
- wherein a ratio of air pressure across a blade of the fan is configured to be less than about 1.50 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition,
- wherein a ratio of a volume of air passing into the bypass duct divided by a volume of air passing into the engine core is greater than about 8.0.
Dependencies are inferred from the claim text.