Gas turbine engine with low stage count low pressure turbine and engine mounting arrangement
A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a high spool, and a low spool including a low pressure turbine that drives the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.
1. A turbofan gas turbine engine comprising:
a fan section including a fan rotor with a plurality of fan blades;
an epicyclic gear train defined about an engine axis of rotation and having a gear reduction ratio greater than 2.5:1;
a low spool including a low pressure compressor and a three or four stage low pressure turbine, wherein the low pressure turbine drives the fan rotor through the gear train;
a high spool including a high pressure compressor and a two stage high pressure turbine, wherein the low and high spools are rotatable about the engine axis of rotation;
a combustor arranged between the high pressure compressor and the high pressure turbine;
a first nacelle which at least partially surrounds a second nacelle and the fan rotor, wherein the second nacelle houses the low spool and the high spool, the first nacelle has a first exit, the second nacelle has a second exit axially aft of the first exit, and the fan section communicates airflow into the first nacelle and the second nacelle and provides an engine bypass ratio greater than 10:1; and
a static structure comprising a fan case and a first static structure component located forward of a second static structure component, the fan case surrounding the fan blades, the first static structure component having a first engine mount location and the second static structure component having a second engine mount location, each of the first engine mount location and the second engine mount location supporting an engine mount when the engine is mounted, wherein the first static structure component is an intermediate case that at least partially surrounds the gear train, and the first engine mount location is axially near the gear train relative to the engine axis of rotation.
2. The turbofan gas turbine engine as recited in claim 1 , wherein the static structure supports a bearing system upon which the low spool, high spool and the fan rotor rotate in operation.
3. The turbofan gas turbine engine as recited in claim 2 , wherein the static structure includes
a high pressure compressor case, a high pressure turbine case, a thrust case, a low pressure turbine case and a turbine exhaust case, wherein the high pressure compressor case is secured to the fan case at the intermediate case, and wherein the second static structure component is a mid-turbine frame within the thrust case.
4. The turbofan gas turbine engine as recited in claim 3 , wherein the mid-turbine frame supports first and second bearings of the bearing system which rotatably support the low spool and the high spool, respectively.
5. The turbofan gas turbine engine as recited in claim 4 , further comprising a low fan pressure ratio less than 1.45 across the fan blades alone, wherein the fan section has a low corrected fan tip speed less than 1150 ft/second, and the low pressure turbine has a pressure ratio greater than 5:1.
6. The turbofan gas turbine engine as recited in claim 5 , wherein the epicyclic gear train is a planetary gear train.
7. The turbofan gas turbine engine as recited in claim 4 , wherein the second engine mount location is not connected to the first static structure component by a thrust link when the engine is mounted.
8. The turbofan gas turbine engine as recited in claim 7 , wherein the second engine mount location reacts an engine thrust in operation.
9. The turbofan gas turbine engine as recited in claim 8 , wherein the intermediate case includes a multiple of circumferentially spaced radially extending struts.
10. The turbofan gas turbine engine as recited in claim 9 , wherein the engine mount includes a rear mount that is attached through the thrust case to the mid-turbine frame when the engine is mounted.
11. The turbofan gas turbine engine as recited in claim 9 , wherein the low pressure turbine has a pressure ratio greater than 5:1.
12. The turbofan gas turbine engine as recited in claim 11 , wherein the epicyclic gear train is a planetary gear train.
13. The turbofan gas turbine engine as recited in claim 12 , wherein the low pressure compressor is a two or four stage compressor.
14. The turbofan gas turbine engine as recited in claim 12 , wherein the bearing system includes third and fourth bearings which rotatably support the low spool and the high spool, respectively, and each of the third and fourth bearings being axially located within the intermediate case with respect to the engine axis of rotation.
15. The turbofan gas turbine engine as recited in claim 14 , wherein the low pressure compressor is a four stage compressor, and the low pressure turbine is a four stage turbine.
16. The turbofan gas turbine engine as recited in claim 15 , wherein the first and second bearings straddle radially extending structural struts which are preloaded in tension.
17. The turbofan gas turbine engine as recited in claim 12 , further comprising a low fan pressure ratio less than 1.45 across the fan blades alone.
18. The turbofan gas turbine engine as recited in claim 17 , wherein the fan section has a low corrected fan tip speed less than 1150 ft/second.
19. The turbofan gas turbine engine as recited in claim 18 , wherein the fan blades have a design angle of incidence, and further comprising:
a fan variable area nozzle axially movable relative to the first nacelle to vary a fan nozzle exit area; and
a controller that controls the fan variable area nozzle to vary the fan nozzle exit area to reduce a fan instability, and to maintain an angle of incidence of the fan blades close to the design angle of incidence at a plurality of flight conditions.
20. The turbofan gas turbine engine as recited in claim 19 , wherein:
the controller reduces the fan nozzle exit area at a cruise flight condition; and
the fan blades have a fixed stagger angle.
21. A turbofan gas turbine engine comprising:
a fan section including a fan rotor with a plurality of fan blades;
a planetary gear train defined about an engine axis of rotation;
a first nacelle which at least partially surrounds a second nacelle and the fan rotor, the first nacelle having a first exit, and the fan section communicates airflow into the first nacelle and the second nacelle and provides an engine bypass ratio greater than 10:1;
a high pressure compressor and a low pressure compressor having two stages;
a three to six stage low pressure turbine and a two stage high pressure turbine, the low pressure turbine driving the fan rotor through the gear train; and
a static structure comprising a first case located forward of a second case, the first case having a first engine mount location and the second case having a second engine mount location, each of the first engine mount location and second engine mount location supporting an engine mount when the engine is mounted, and wherein the first engine mount location is axially near the gear train.
22. The turbofan gas turbine engine as recited in claim 21 , wherein the static structure supports a bearing system upon which the low pressure turbine, high pressure turbine and fan rotor rotate in operation.
23. The turbofan gas turbine engine as recited in claim 22 , wherein the gear train has a gear reduction ratio greater than 2.5:1.
24. The turbofan gas turbine engine as recited in claim 23 , wherein the low pressure turbine has a pressure ratio greater than 5:1.
25. The turbofan gas turbine engine as recited in claim 24 , wherein the low pressure turbine is a three or four stage turbine.
26. The turbofan gas turbine engine as recited in claim 25 , wherein the second engine mount location is not connected to the first static structure component by a thrust link when the engine is mounted.
27. The turbofan gas turbine engine as recited in claim 26 , further comprising a low fan pressure ratio less than 1.45 across the fan blades alone.
28. The turbofan gas turbine engine as recited in claim 27 , wherein:
the static structure includes a fan case, an intermediate case and a high pressure compressor case; and
the first case is the intermediate case, and the high pressure compressor case is secured to the fan case at the intermediate case.
29. The turbofan gas turbine engine as recited in claim 27 , wherein:
the static structure includes a high pressure turbine case, a thrust case, a low pressure turbine case and a turbine exhaust case;
the second case is the thrust case, and a mid-turbine frame within the thrust case defines the second engine mount location and supports first and second bearings of the bearing system; and
the second engine mount location reacts an engine thrust in operation.
30. The turbofan gas turbine engine as recited in claim 29 , wherein:
the static structure includes a fan case, an intermediate case and a high pressure compressor case; and
the first case is the intermediate case, the high pressure compressor case is secured to the fan case at the intermediate case, and the intermediate case includes a multiple of circumferentially spaced radially extending struts.