GEARED GAS TURBINE ENGINE ARCHITECTURE FOR ENHANCED EFFICIENCY

• US 20150027101A1
• Filed: 10/03/2013
• Published: 01/29/2015
• Est. Priority Date: 01/21/2013
• Status: Abandoned Application

• Current Assignee: Raytheon Technologies Corporation d/b/a RTX (Rtx Corporation)
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Hasel, Karl L.
• Application Number:
  US14/044,971
• Publication Number:
  US 20150027101A1
• Time in Patent Office: Days
• US Class Current: 60/226.1
• CPC Class Codes: F02C 7/36 Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user F02C3/107 - F02C3/13 and F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F05D 2260/4031 as in toothed gearing

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.

Claims

1. 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.

2. The gas turbine engine of claim 1, wherein the first compressor section is a high pressure compressor.

3. The gas turbine engine of claim 1, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when cruising at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

4. The gas turbine engine of claim 1, wherein a ratio of air pressure across a blade of the fan is configured to be less than about 1.45 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

5. The gas turbine engine of claim 1, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when the fan stream exhaust velocity is less than about 1175 feet per second.

6. The gas turbine engine of claim 1, wherein the fan has less than about 26 fan blades.

7. The gas turbine engine of claim 6, wherein the fan has less than about 20 fan blades.

8. 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.

9. The gas turbine engine of claim 8, wherein the first compressor section is a high pressure compressor.

10. The gas turbine engine of claim 8, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when cruising at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

11. The gas turbine engine of claim 8, wherein a ratio of air pressure across a blade of the fan is configured to be less than about 1.45 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

12. The gas turbine engine of claim 8, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when the fan stream exhaust velocity is less than about 1175 feet per second.

13. The gas turbine engine of claim 8, wherein the fan has less than about 26 fan blades.

14. The gas turbine engine of claim 13, wherein the fan has less than about 20 fan blades.

15. 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.

16. The gas turbine engine of claim 15, wherein the first compressor section is a high pressure compressor.

17. The gas turbine engine of claim 15, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when cruising at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

18. The gas turbine engine of claim 15, wherein a ratio of air pressure across a blade of the fan for the engine is configured to be less than about 1.45 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

19. The gas turbine engine of claim 15, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when the fan stream exhaust velocity is less than about 1175 feet per second.

20. 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.

21. The gas turbine engine of claim 20, wherein the ratio of air pressure across a blade of the fan is configured to be less than about 1.45 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

22. The gas turbine engine of claim 20, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when the fan stream exhaust velocity is less than about 1175 feet per second.

23. The gas turbine engine of claim 21, wherein the engine is configured so that the ratio of the fan stream exhaust velocity divided by the primary stream exhaust velocity is in the range when the fan stream exhaust velocity is less than about 1175 feet per second.

24. The gas turbine engine of claim 21, wherein the engine is configured so that a tip speed of the fan blade is less than about 1200 ft/second at about 35.000 feet and when operating at about the 0.80 Mach number cruise power condition.

25. The gas turbine engine of claim 24, wherein the engine is configured so that a tip speed of the fan blade is less than about 1150 ft/second at about 35.000 feet and when operating at about the 0.80 Mach number cruise power condition.

26. The gas turbine engine of claim 23, wherein the engine is configured so that a tip speed of the fan blade is less than about 1200 ft/second at about 35,000 feet and when operating at about the 0.80 Mach number cruise power condition.

27. The gas turbine engine of claim 26, wherein the engine is configured so that a tip speed of the fan blade is less than about 1150 ft/second at about 35,000 feet and when operating at about the 0.80 Mach number cruise power condition.

28. The gas turbine engine of claim 20, further comprising:

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,

29. The gas turbine engine of claim 28, wherein the ratio of air pressure across a blade of the fan is configured to be less than about 1.45 at about 35,000 feet and when operating at about a 0.80 Mach number cruise power condition.

30. The gas turbine engine of claim 29, wherein the engine is configured so that a tip speed of the fan blade is less than about 1200 ft/second at about 35.000 feet and when operating at about the 0.80 Mach number cruise power condition.