Geared turbofan gas turbine engine architecture

Patent metadata

Field	Value
Publication number	US-9695751-B2
Application number	US-201615395277-A
Country	US
Kind code	B2
Filing date	Dec 30, 2016
Priority date	Jan 31, 2012
Publication date	Jul 4, 2017
Grant date	Jul 4, 2017

Abstract

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a fan including a plurality of fan blades rotatable about an axis;
a compressor section;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, the turbine section including a fan drive turbine and a second turbine, wherein the second turbine is disposed forward of the fan drive turbine and the fan drive turbine includes a plurality of turbine rotors with a ratio between a number of fan blades and a number of fan drive turbine rotors is between 2.5 and 8.5; and
a speed change system configured to be driven by the fan drive turbine to rotate the fan about the axis; and
a power density at Sea Level Takeoff greater than or equal to 1.5 lbf/in 3 and less than or equal to 5.5 lbf/in 3 and defined as thrust in lbf measured by a volume of the turbine section in in 3 measured between an inlet of a first turbine vane in said second turbine to an exit of a last rotating airfoil stage in said fan drive turbine.

2. The gas turbine engine as recited in claim 1 , wherein the fan drive turbine has from three to six stages.

3. The gas turbine engine as recited in claim 2 , wherein said number of fan blades is less than 18 and the second turbine has two stages.

4. The gas turbine engine as recited in claim 3 , further comprising a frame structure positioned between the fan drive turbine and the second turbine, and a plurality of vanes associated with the frame structure, and a flow path through said frame structure being part of the volume of the turbine section.

5. The gas turbine engine as recited in claim 3 , wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, the first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is greater than or equal to 0.5 and less than or equal to 1.5.

6. The gas turbine engine as recited in claim 5 , wherein the performance ratio is above or equal to about 0.8.

7. The gas turbine engine as recited in claim 6 , wherein the performance ratio is above or equal to about 1.0.

8. The gas turbine engine as recited in claim 7 , wherein the power density is greater than or equal to 3.0 lbf/in 3 .

9. The gas turbine engine as recited in claim 1 , wherein said number of fan blades is less than 18 and the second turbine has two stages.

10. The gas turbine engine as recited in claim 9 , further comprising a frame structure positioned between the fan drive turbine and the second turbine, and a plurality of vanes associated with the frame structure, and a flow path through said frame structure being part of the volume of the turbine section.

11. The gas turbine engine as recited in claim 10 , wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, the first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is greater than or equal to 0.5 and less than or equal to 1.5.

12. The gas turbine engine as recited in claim 11 , wherein the performance ratio is above or equal to about 0.8.

13. The gas turbine engine as recited in claim 9 , wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, the first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is greater than or equal to about 0.8 and less than or equal to 1.5.

14. The gas turbine engine as recited in claim 1 , wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, the first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is greater than or equal to about 0.8 and less than or equal to 1.5.

15. A gas turbine engine comprising:
a fan including a plurality of fan blades rotatable about an axis;
a compressor section;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, the turbine section including a fan drive turbine and a second turbine, wherein the second turbine is disposed forward of the fan drive turbine and the fan drive turbine includes a plurality of turbine rotors with a ratio between a number of fan blades and a number of fan drive turbine rotors is between 2.5 and 8.5; and
a speed change system having a gear reduction, the speed change system configured to be driven by the fan drive turbine to rotate the fan about the axis; and
a power density at Sea Level Takeoff greater than or equal to 1.5 and less than or equal to 5.5 lbf/in 3 and defined as thrust in lbf divided by a volume of the turbine section in inch 3 measured between an inlet of a first turbine vane in said second turbine to an exit of a last rotating airfoil stage in said fan drive turbine.

16. The gas turbine engine as recited in claim 15 , wherein said number of fan blades is less than 18, and the second turbine has two stages.

17. The gas turbine engine as recited in claim 15 , wherein the fan drive turbine has a first exit area and rotates at a first speed, the second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed, the first and second speeds being redline speeds, a first performance quantity is defined as the product of the first speed squared and the first area, a second performance quantity is defined as the product of the second speed squared and the second area, and a performance ratio of the first performance quantity to the second performance quantity is greater than or equal to about 0.8 and less than or equal to 1.5.

18. The gas turbine engine as recited in claim 17 , wherein said number of fan blades is less than 18 and the second turbine has two stages.

19. The gas turbine engine as recited in claim 18 , wherein the performance ratio is above or equal to about 1.0.

20. The gas turbine engine as recited in claim 19 , further comprising a frame structure positioned between the fan drive turbine and the second turbine, and a plurality of vanes associated with the frame structure, and a flow path through said frame s

Assignees

• United Technologies Corp

Inventors

• Kupratis Daniel Bernard
• Schwarz Frederick M

Classifications

• F02C7/36 Primary

  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 title

• F01D9/02

  Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} · CPC title

• F04D27/009

  by bleeding, by passing or recycling fluid · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F05D2240/60

  Shafts · CPC title