Geared gas turbine engine arrangement with core split power ratio

1. A gas turbine engine comprising:

a propulsor section including a propulsor having a plurality of blades, the propulsor having a low corrected tip speed of less than 1150 ft/sec, and the low corrected tip speed being an actual tip speed divided by [(Tram ° R)/(518.7° R)]{circumflex over ( )}0.5;

a gear arrangement including a gear reduction ratio greater than 2.6;

a compressor section including a low pressure compressor and a high pressure compressor, and the high pressure compressor including greater than or equal to 8 stages;

a turbine section including a low pressure turbine and a high pressure turbine, the high pressure turbine driving the high pressure compressor, the low pressure turbine driving the low pressure compressor and driving the gear arrangement to drive the propulsor, the low pressure turbine including three or more stages, and the low pressure compressor including a greater number of stages than the high pressure turbine;

wherein the low pressure turbine includes an inlet, an outlet, and a low pressure turbine pressure ratio greater than 5, and the low pressure turbine pressure ratio is pressure measured prior to the inlet as related to pressure at the outlet prior to any exhaust nozzle; and

wherein a core split power ratio is provided by a second power input to the high pressure compressor measured in horsepower divided by a first power input to the low pressure compressor measured in horsepower, the core split power ratio being equal to, or greater than, 1.0 and less than, or equal to, 1.4.

2. The gas turbine engine as set forth in claim 1 , wherein the low pressure compressor includes 3 or more stages.

3. The gas turbine engine as set forth in claim 2 , wherein the high pressure turbine includes two stages.

4. The gas turbine engine as set forth in claim 3 , wherein:

the gas turbine engine is a two-spool engine including a low spool and a high spool;

the low spool includes the low pressure compressor, the low pressure turbine and an inner shaft interconnecting an input of the gear arrangement and the low pressure turbine; and

the high spool includes the high pressure compressor, the high pressure turbine and an outer shaft interconnecting the high pressure compressor and the high pressure turbine.

5. The gas turbine engine as set forth in claim 4 , wherein the low pressure turbine includes six or fewer stages.

6. The gas turbine engine as set forth in claim 5 , wherein the propulsor and the low pressure turbine are rotatable in opposite directions via the gear arrangement.

7. The gas turbine engine as set forth in claim 5 , wherein the propulsor and the low pressure turbine are rotatable in the same direction via the gear arrangement.

8. The gas turbine engine as set forth in claim 5 , wherein the gear arrangement includes an epicyclic gear train.

9. The gas turbine engine as set forth in claim 8 , wherein the core split power ratio is less than 1.3.

10. The gas turbine engine as set forth in claim 9 , wherein the gear arrangement has a thermal efficiency at sea-level takeoff of greater than 98.7% as measured by an oil temperature rise between an inlet and an outlet of the gear arrangement.

11. The gas turbine engine as set forth in claim 10 , wherein the epicyclic gear train is a planetary gear system, and the gear reduction ratio is at least 2.9.

12. The gas turbine engine as set forth in claim 9 , wherein the gear reduction ratio is at least 2.9.

13. The gas turbine engine as set forth in claim 12 , wherein the low pressure compressor section includes 4 or more stages.

14. The gas turbine engine as set forth in claim 12 , wherein the core split power ratio is less than 1.27.

15. The gas turbine engine as set forth in claim 14 , wherein the gear arrangement has a thermal efficiency at sea-level takeoff of greater than 98.7% as measured by an oil temperature rise between an inlet and an outlet of the gear arrangement.

16. The gas turbine engine as set forth in claim 14 , wherein the propulsor section is a fan section, the propulsor is a fan, and further comprising a fan case surrounding the plurality of blades to define a bypass duct, wherein the fan section delivers a portion of air into the compressor section, and a portion of air into the bypass duct, and wherein 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.

17. The gas turbine engine as set forth in claim 16 , wherein the fan and the low pressure turbine are rotatable in opposite directions via the gear arrangement.

18. The gas turbine engine as set forth in claim 16 , wherein the epicyclic gear train is a planetary gear system.

19. The gas turbine engine as set forth in claim 18 , wherein the fan section defines a fan pressure ratio of less than 1.45 across the blade alone at cruise at 0.8 Mach and 35,000 feet.

20. The gas turbine engine as set forth in claim 19 , wherein an overall pressure ratio is provided by the combination of the low pressure compressor, the high pressure compressor and a fan root pressure rise of the fan section, and the overall pressure ratio is equal to or greater than 36.

21. The gas turbine engine as set forth in claim 20 , wherein the turbine section includes a mid-turbine frame between the high pressure turbine and the low pressure turbine, the mid-turbine frame includes airfoils in a core flow path, and the mid-turbine frame supports bearing systems in the turbine section.

22. The gas turbine engine as set forth in claim 21 , wherein the overall pressure ratio is equal to or greater than 50.

23. The gas turbine engine as set forth in claim 22 , wherein the fan pressure ratio is less than 1.40 across the blade alone at cruise at 0.8 Mach and 35,000 feet.

24. The gas turbine engine as set forth in claim 23 , wherein the low pressure compressor is a three-stage compressor.

25. The gas turbine engine as set forth in claim 23 , wherein the low pressure compressor is a four-stage compressor.

26. The gas turbine engine as set forth in claim 23 , wherein the gear arrangement has a thermal efficiency at sea-level takeoff of greater than 98.7% as measured by an oil temperature rise between an inlet and an outlet of the gear arrangement.

27. The gas turbine engine as set forth in claim 26 , wherein the bypass ratio is greater than 12 at cruise at 0.8 Mach and 35,000 feet.

28. The gas turbine engine as set forth in claim 27 , wherein:

the fan section has a blade efficiency greater than 94.5%, the blade efficiency defined as an amount of thrust generated by the blade compared to the work provided to the blade by the gear arrangement;

the low pressure turbine has a thermal efficiency greater than 90.9%, the thermal efficiency of the low pressure turbine defined as a pressure decrease between the inlet and the outlet as compared to the work transferred to the low spool; and

the low pressure compressor has a thermal efficiency of at least 87% and delivers air to the high pressure compressor having a thermal efficiency of at least 85.1%, the thermal efficiency of the high and low pressure compressors defined as a pressure rise versus a temperature rise between an inlet and an outlet of a respective one of the low pressure and high pressure compressors.

29. The gas turbine engine as set forth in claim 28 , wherein the overall pressure ratio is less than or equal to 80.

30. The gas turbine engine as set forth in claim 29 , wherein the core split power ratio is no more than 1.1.