Gas turbine engine with front support arrangement
A gas turbine engine includes a propulsor having blades extending from a propulsor hub. A compressor section includes a first compressor and a second compressor. A speed reduction device includes an epicyclic gear system. The epicyclic gear system includes a sun gear, a plurality of intermediate gears, a carrier supporting the intermediate gears and the propulsor hub, and a ring gear. A turbine section includes a first turbine and a second turbine. The second turbine drives the propulsor through the epicyclic gear system. The epicyclic gear system is straddled by forward and aft bearings that engage the carrier.
1. A gas turbine engine comprising:
a propulsor section including a propulsor having blades extending from a propulsor hub;
a propulsor shaft that drives the propulsor, the propulsor rotatable about an engine longitudinal axis;
a compressor section including a compressor inlet having an entrance at a splitter, a first compressor and a second compressor, wherein the blades of the propulsor are forward of the splitter relative to the engine longitudinal axis such that the blades deliver airflow to the entrance of the compressor inlet in response to rotation of the blades;
a speed reduction device including an epicyclic gear system, the epicyclic gear system including a gear reduction ratio greater than 2.3, a sun gear, a plurality of intermediate gears, a carrier supporting the intermediate gears and the propulsor hub, and a ring gear;
a turbine section including a first turbine and a second turbine, wherein the second turbine drives the propulsor through the epicyclic gear system;
wherein the epicyclic gear system is straddled by forward and aft bearings that engage the carrier; and
wherein the speed reduction device turns the propulsor and the first compressor at a common rotational speed.
2. The gas turbine engine as recited in claim 1 , wherein the propulsor section is a fan section, the propulsor is a fan, the propulsor hub is a fan hub, and an outer housing surrounds the fan to define a bypass duct.
3. The gas turbine engine as recited in claim 2 , wherein:
the fan section delivers a portion of air into the compressor section, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is greater than 10 at cruise at 0.8 Mach and 35,000 feet.
4. The gas turbine engine as recited in claim 1 , wherein:
the gas turbine engine has only one propulsor stage, and the propulsor stage comprises the blades of the propulsor.
5. The gas turbine engine as recited in claim 4 , wherein:
the first compressor includes three stages;
the carrier is fixed relative to an engine static structure; and
the ring gear establishes an output of the speed reduction device.
6. The gas turbine engine as recited in claim 1 , wherein:
the forward bearing is axially aligned with the blades of the propulsor relative to the engine longitudinal axis.
7. A gas turbine engine comprising:
a propulsor section including a propulsor having blades extending from a propulsor hub;
a propulsor shaft that drives the propulsor, the propulsor rotatable about an engine longitudinal axis;
a compressor section including a first compressor and a second compressor;
a speed reduction device including an epicyclic gear system, the epicyclic gear system including a gear reduction ratio greater than 2.3, a sun gear, a plurality of intermediate gears, a carrier supporting the intermediate gears and the propulsor hub, and a ring gear;
a turbine section including a first turbine and a second turbine, wherein the second turbine drives the propulsor through the epicyclic gear system;
wherein the epicyclic gear system is straddled by forward and aft bearings that engage the carrier; and
wherein the forward bearing supports radial and thrust loads from the propulsor hub, and the aft bearing carries radial loads from the propulsor.
8. The gas turbine engine as recited in claim 7 , wherein the speed reduction device turns the propulsor and the first compressor at a common rotational speed.
9. The gas turbine engine as recited in claim 8 , wherein the first compressor includes a plurality of stages.
10. The gas turbine engine as recited in claim 9 , wherein the second compressor includes a compression ratio of at least 20:1.
11. The gas turbine engine as recited in claim 9 , wherein:
the carrier is fixed relative to an engine static structure;
the ring gear drives the propulsor hub; and
the forward and aft bearings support rotation of the propulsor hub.
12. The gas turbine engine as recited in claim 11 , wherein the epicyclic gear system is radially inward and axially aligned with the first compressor relative to the engine longitudinal axis.
13. The gas turbine engine as recited in claim 11 , wherein the second turbine includes three stages.
14. The gas turbine engine as recited in claim 11 , wherein the first turbine includes two stages and drives the second compressor.
15. The gas turbine engine as recited in claim 14 , wherein the second turbine includes an inlet, an outlet and a turbine pressure ratio of greater than 5:1, and the turbine pressure ratio is pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle.
16. The gas turbine engine as recited in claim 15 , wherein the first compressor includes three stages.
17. The gas turbine engine as recited in claim 16 , wherein the first compressor includes a compression ratio of at least 2:1.
18. The gas turbine engine as recited in claim 17 , wherein the second compressor includes a compression ratio of at least 20:1.
19. The gas turbine engine as recited in claim 16 , wherein the second compressor includes a compression ratio of at least 20:1.
20. The gas turbine engine as recited in claim 19 , wherein the first compressor includes a compression ratio of at least 2:1.
21. The gas turbine engine as recited in claim 11 , wherein:
the ring gear drives the propulsor hub and the first compressor, and the first compressor abuts the propulsor hub at an interface.
22. The gas turbine engine as recited in claim 11 , wherein the forward bearing supports rotation of the propulsor hub, and the forward bearing is axially aligned with the propulsor hub relative to the engine longitudinal axis.
23. The gas turbine engine as recited in claim 7 , wherein the propulsor section is a fan section, the propulsor is a fan, the propulsor hub is a fan hub, and an outer housing surrounds the fan to define a bypass duct.
24. The gas turbine engine as recited in claim 23 , wherein:
the fan section delivers a portion of air into the compressor section, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is greater than 10 at cruise at 0.8 Mach and 35,000 feet.
25. The gas turbine engine as recited in claim 24 , wherein the speed reduction device turns the propulsor and the first compressor at a common rotational speed.
26. The gas turbine engine as recited in claim 25 , wherein the fan has a fan pressure ratio of less than 1.45 across the blades alone at cruise at 0.8 Mach and 35,000 feet.
27. The gas turbine engine as recited in claim 26 , wherein the first compressor includes a plurality of stages.
28. The gas turbine engine as recited in claim 27 , wherein the first turbine includes two stages and drives the second compressor, and the second turbine includes three stages.
29. The gas turbine engine as recited in claim 27 , wherein:
the carrier is fixed relative to an engine static structure;
the ring gear drives the fan hub; and
the forward and aft bearings support rotation of the fan hub.
30. The gas turbine engine as recited in claim 27 , wherein the forward bearing supports rotation of the fan hub, and the forward bearing is axially aligned with the fan hub relative to the engine longitudinal axis.
31. The gas turbine engine as recited in claim 30 , wherein:
the first compressor includes three stages, the second compressor includes a greater number of stages than the first compressor;
the first turbine includes two stages and drives the second compressor, and the second turbine includes a greater number of stages than the first turbine; and
the turbine section includes a mid-turbine frame arranged between the first turbine and the second turbine with respect to the engine longitudinal axis, the mid-turbine frame supports a bearing, and the mid-turbine frame includes airfoils in a core airflow path.
32. The gas turbine engine as recited in claim 24 , wherein the first compressor includes a greater number of stages than the first turbine.
33. The gas turbine engine as recited in claim 32 , wherein the first compressor includes a compression ratio of at least 2:1, and the second compressor includes a compression ratio of at least 20:1.