Geared turbofan arrangement with core split power ratio
A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including at least a first compressor section and a second compressor section. A power ratio is provided by the combination of the first compressor section and the second compressor section. A method of design a gas turbine engine is also disclosed.
1. A gas turbine engine comprising:
a fan section including a plurality of fan blades;
a compressor section, including at least a first compressor section and a second compressor section;
a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least an epicyclic gear arrangement to drive said fan section and said first compressor section, said turbine section including at least two turbine stages upstream of said fan drive turbine; and
wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said 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 said fan drive turbine includes six or fewer stages.
3. The gas turbine engine as set forth in claim 2 , wherein said power ratio is less than 1.27.
4. The gas turbine engine as set forth in claim 2 , wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 10.
5. The gas turbine engine as set forth in claim 2 , wherein a gear ratio of said gear arrangement is greater than 2.6.
6. The gas turbine engine as set forth in claim 2 , wherein an overall pressure ratio being provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36.
7. The gas turbine engine as set forth in claim 1 , wherein said fan section defines a fan pressure ratio less than 1.50 across the fan blades alone, said first turbine section is configured to rotate at least 2.6 times faster than said fan section, and an overall pressure ratio being provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36.
8. The gas turbine engine as set forth in claim 1 , wherein said first compressor section includes 3 or more stages and said second compressor section includes between 8 and 15 stages.
9. The gas turbine engine as recited in claim 8 , wherein said fan drive turbine includes three 3 to six 6 stages.
10. The gas turbine engine as recited in claim 9 , wherein said fan drive turbine defines a fan drive turbine pressure ratio that is greater than five.
11. The gas turbine engine as set forth in claim 7 , wherein said power ratio is less than 1.27, said overall pressure ratio is equal to or greater than 50, and said gear arrangement is a planetary gear system.
12. A gas turbine engine comprising:
a fan section including a plurality of fan blades;
a compressor section, including at least a first compressor section and a second compressor section, said first compressor section including 3 or more stages and said second compressor section including 6 or more stages;
a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least a gear arrangement to drive said fan section, said turbine section including at least 2 turbine stages upstream of said fan drive turbine; and
wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio being equal to or greater than 1.0; and
wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to, or greater than, 36.
13. The gas turbine engine as set forth in claim 12 , wherein said fan section defines a fan pressure ratio less than 1.50 across the fan blades alone, a gear ratio of said gear arrangement is greater than 2.6, said turbine section includes at least 2 turbine stages upstream of said fan drive turbine, and said fan drive turbine includes 3 to 6 stages.
14. The gas turbine engine as set forth in claim 12 , wherein said power ratio is between 1.0 and 1.4.
15. The gas turbine engine as set forth in claim 12 , wherein said fan section defines a fan pressure ratio less than 1.45 across the fan blades alone.
16. The gas turbine engine as recited in claim 15 , wherein said fan drive turbine defines a fan drive turbine pressure ratio that is greater than five (5).
17. The gas turbine engine as set forth in claim 16 , wherein said power ratio is less than 1.27.
18. The gas turbine engine as recited in claim 17 , wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 12.
19. The gas turbine engine as set forth in claim 13 , wherein said gear arrangement is a planetary gear system, and said overall pressure ratio is equal to or greater than 60.
20. A gas turbine engine comprising:
a fan section;
a compressor section, including at least a first compressor section and a second compressor section;
a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive at least a gear arrangement to drive said fan section;
wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio being equal to or less than 1.4; and
wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36.
21. The gas turbine engine as set forth in claim 20 , wherein said first compressor section includes 3 or more stages, said second compressor section includes 6 or more stages, said turbine section includes at least 2 turbine stages upstream of said fan drive turbine, and said fan drive turbine includes three 3 to six 6 stages.
22. The gas turbine engine as set forth in claim 20 , wherein said power ratio is greater than or equal to 1.0.
23. The gas turbine engine as set forth in claim 20 , wherein said fan section is configured to deliver a portion of air into said compressor section, and a portion of air into a bypass duct, and wherein a bypass ratio, which is defined as a volume of air passing to said bypass duct compared to a volume of air passing into said compressor section, is equal to or greater than 10.
24. The gas turbine engine as set forth in claim 20 , wherein said overall pressure ratio is equal to or greater than 60.
25. A method of designing a gas turbine engine comprising:
providing a fan section;
providing a compressor section, including a first compressor section and a second compressor section;
providing a turbine section including at least one turbine to drive said second compressor section and a fan drive turbine to drive said fan section via a gear arrangement;
wherein a power ratio is provided by the combination of said first compressor section and said second compressor section, with said power ratio being provided by a second power input to said second compressor section divided by a first power input to said first compressor section, said power ratio is equal to or less than 1.4 at a predetermined design target; and
wherein an overall pressure ratio is provided by the combination of said first compressor section, said second compressor section and a fan root pressure rise of said fan section, said overall pressure ratio being equal to or greater than 36 at the predetermined design target.
26. The method as set forth in claim 25 , wherein said predetermined design target is defined at sea level and at a static, full-rated takeoff power condition.
27. The method as set forth in claim 25 , wherein said predetermined design target is defined at a cruise condition.
28. The method as set forth in claim 25 , wherein said power ratio is greater than or equal to 1.0.
29. The method as set forth in claim 25 , wherein said first compressor section includes 3 or more stages, said second compressor section includes 6 or more stages, said second compressor section is positioned downstream of said first compressor section, said fan drive turbine includes three 3 to six 6 stages, and said at least one turbine includes fewer stages than said fan drive turbine.
30. The method as set forth in claim 29 , wherein said power ratio is less than 1.27, and said overall pressure ratio is equal to or greater than 60.