Method for setting a gear ratio of a fan drive gear system of a gas turbine engine
A gas turbine engine includes a fan section including a fan that is rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0.
1. A gas turbine engine comprising:
a fan section including a fan rotatable about an axis;
a speed reduction device connected to the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5, wherein a bypass ratio is greater than about 11.0.
2. The gas turbine engine of claim 1 , wherein the speed reduction device includes a star gear system gear ratio of at least 2.6.
3. The gas turbine engine of claim 2 , wherein the speed reduction device includes a system gear ratio less than or equal to 4.1.
4. The gas turbine engine of claim 3 , including the bypass ratio is less than about 22.0.
5. The gas turbine engine of claim 4 , wherein the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.
6. The gas turbine engine of claim 1 , wherein the star system includes a sun gear, a plurality of star gears, a ring gear, and a carrier.
7. The gas turbine engine of claim 6 , wherein each of the plurality of star gears includes at least one bearing.
8. The gas turbine engine of claim 7 , wherein the carrier is fixed from rotation.
9. The gas turbine engine of claim 8 , wherein a low pressure turbine is mechanically attached to the sun gear.
10. The gas turbine engine of claim 9 , wherein a fan section is mechanically attached to the ring gear.
11. The gas turbine engine of claim 1 , wherein an input of the speed reduction device is rotatable in a first direction and an output of the speed reduction device is rotatable in a second direction opposite to the first direction.
12. The gas turbine engine of claim 11 , including a low pressure turbine section in communication with the speed reduction device, wherein the low pressure turbine section includes at least three stages.
13. A method of improving performance of a gas turbine engine comprising:
determining fan tip speed boundary conditions for at least one fan blade of a fan section;
determining rotor boundary conditions for a rotor of a low pressure turbine;
utilizing stress level constraints in the rotor of the low pressure turbine and the at least one fan blade to determine if the rotary speed of the fan section and the low pressure turbine will meet a desired number of operating cycles, wherein a bypass ratio is greater than about 6.0, wherein a speed reduction device connects the fan section and the low pressure turbine and includes a star gear ratio of at least about 1.5.
14. The method of claim 13 , wherein the star gear ratio is less than about 4.1.
15. The method of claim 14 , wherein a fan pressure ratio is below 1.7.
16. The method of claim 15 , wherein a fan pressure ratio is below 1.48.
17. The method of claim 16 , wherein a bypass ratio is between about 11 and about 22.
18. The method of claim 17 , wherein a fan blade tip speed of the at least one fan blade is less than 1400 fps.
19. The method of claim 13 , wherein if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and the number of stages of the low pressure turbine is increased.
20. The method of claim 19 , wherein if a stress level in the rotor or the at least one fan blade is too high to meet a desired number of operating cycles, a gear ratio of a gear reduction device is lowered and an annular area of the low pressure turbine is increased.
21. A fan drive gear module for a gas turbine engine comprising:
a star drive gear system with a speed reduction ratio of at least 1.5, wherein the star drive gear system is configured to drive a gas turbine engine with a bypass ratio greater than about 11.0.
22. The fan drive gear module of claim 21 , wherein the speed reduction ratio is greater than about 2.6 and less than or equal to 4.1.
23. The fan drive gear module of claim 21 , wherein the star drive gear system is configured to drive a fan section with a fan blade tip speed greater than about 1000 ft/sec and less than about 1400 ft/sec.
24. The fan drive gear module of claim 21 , wherein the bypass ratio is less than about 22.0.
25. A method of designing a gas turbine engine comprising:
selecting fan tip speed boundary conditions for at least one fan blade of a fan section of a gas turbine engine;
selecting rotor boundary conditions for a rotor of a fan drive turbine of the gas turbine engine;
determining stress level constraints in the rotor of the fan drive turbine and the at least one fan blade to determine if the rotary speed of the fan section and the fan drive turbine will meet a desired number of operating cycles, wherein a bypass ratio is greater than about 6.0; and
connecting the fan section and the drive turbine through a speed reduction device that includes a speed reduction ratio of at least about 1.5.
26. The method of claim 25 , wherein a fan pressure ratio is below 1.7.
27. The method of claim 26 , wherein the speed reduction ratio is less than about 4.1.
28. The method of claim 27 , wherein the bypass ratio of the gas turbine engine is between about 11 and about 22.
29. The method of claim 25 , including lowering the speed reduction ratio of the speed reduction device and increasing a number of stages of the fan drive turbine responsive to determining that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.
30. The method of claim 26 , including lowering the speed reduction ratio of the speed reduction device and increasing an annular area of the fan drive turbine responsive to determining that a stress level in the rotor or the at least one fan blade is outside a predefined number of operating cycles.