Geared turbofan engine gearbox arrangement
A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section having a plurality of fan blades. The plurality of fan blades has a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath. A core engine includes a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section. A method of designing a gas turbine engine is also disclosed.
1. A gas turbine engine comprising:
a fan section including a fan having a plurality of fan blades that deliver airflow to a bypass duct, the plurality of fan blades each having a leading edge, a trailing edge, a peak tip radius Rt defined at the leading edge and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath at the leading edge, and a ratio of Rh to Rt being less than 0.4; and
a core engine including a first turbine that drives a first compressor, and a fan drive turbine that drives the fan section, the fan drive turbine including 5 stages, and the fan drive turbine including fewer stages than the first compressor;
a second compressor;
wherein the fan section has only a single fan stage comprising the fan;
wherein the core engine defines a core flowpath and a longitudinal axis;
wherein a second inboard boundary of the core flowpath has a radius R 1 defined at a first stage of the second compressor and has a radius R 2 defined at a splitter rim, each radius R 1 , R 2 are defined relative to the longitudinal axis, and the splitter rim guides flow into the core flowpath; and
wherein a ratio of an axial length L 10 between the radii R 1 , R 2 to the radius R 2 is less than 1.2.
2. The gas turbine engine as recited in claim 1 , wherein the second compressor is upstream of the first compressor, and the second compressor includes a greater number of stages than the first turbine.
3. The gas turbine engine as recited in claim 1 , wherein:
the ratio of R 1 to R 2 is greater than 0.5.
4. The gas turbine engine as recited in claim 3 , wherein the ratio of R 1 to R 2 is between 0.55 and 1.0.
5. The gas turbine engine as recited in claim 3 , wherein the ratio of Rh to Rt is less than 0.30.
6. The gas turbine engine as recited in claim 5 , wherein the first turbine includes two stages.
7. The gas turbine engine as recited in claim 1 , comprising a geared architecture that rotates the fan section at a different speed than the fan drive turbine.
8. The gas turbine engine as recited in claim 6 , wherein:
the second compressor is upstream of the first compressor and is downstream of the fan, the second compressor including a greater number of stages than the first turbine, and the fan drive turbine drives each stage of the second compressor; and
the ratio of R 1 to R 2 is between 0.55 and 1.0.
9. A gas turbine engine comprising:
a fan section including a fan having a plurality of fan blades that deliver airflow to a bypass duct, the plurality of fan blades each having leading edge, a trailing edge, a peak tip radius Rt defined at the leading edge and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath at the leading edge, and a ratio of Rh to Rt being less than 0.4;
a core engine defining a core flowpath and a longitudinal axis, the core engine including a first turbine that drives a first compressor, a second compressor, and a fan drive turbine that drives the fan section and arranged aft of the first turbine;
wherein a second inboard boundary of the core flowpath has a radius R 1 defined at a first stage of the second compressor and a radius R 2 defined at a splitter rim, each radius R 1 , R 2 defined relative to the longitudinal axis, the splitter rim guides flow into the core flowpath, and a ratio of R 1 to R 2 is greater than 0.5;
wherein the tan section has only a single fan stage comprising the fan; and
wherein a ratio of an axial length L 10 between the radii R 1 , R 2 to the radius R 2 is less than 1.2.
10. The gas turbine engine as recited in claim 9 , wherein the ratio of R 1 to R 2 is between 0.55 and 1.0.
11. The gas turbine engine as recited in claim 10 , wherein the ratio of Rh to Rt is less than 0.30.
12. The gas turbine engine as recited in claim 9 , wherein six or more turbine stages are arranged downstream of the first turbine.
13. The gas turbine engine as recited in claim 12 , wherein the first turbine includes two stages.
14. The gas turbine engine as recited in claim 9 , comprising a geared architecture that rotates the fan section at a different speed than the fan drive turbine.
15. The gas turbine engine as recited in claim 11 , wherein:
the second compressor is upstream of the first compressor and is downstream of the fan, the second compressor including a greater number of stages than the first turbine, and the fan drive turbine drives each stage of the second compressor.
16. The gas turbine engine as recited in claim 15 , wherein:
the first turbine includes two blade stages; and
six or more turbine blade stages are arranged downstream of the first turbine.
17. A method of designing a gas turbine engine comprising:
providing a fan section including a fan having a plurality of fan blades that deliver airflow to a bypass duct, the fan blades each having a leading edge, a trailing edge, a peak tip radius Rt defined at the leading edge and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath at the leading edge, and a ratio of Rh to Rt being less than 0.4; and
providing a core engine including a first turbine that drives a first compressor, a second compressor, and a fan drive turbine that drives the fan section, the fan drive turbine including 5 stages, the first compressor including a greater number of stages than the fan drive turbine, and the second compressor including a greater number of stages than the first turbine;
wherein the fan section has only a single fan stage comprising the fan;
wherein the core engine defines a core flowpath and a longitudinal axis;
wherein a second inboard boundary of the core flowpath has a radius R 1 defined at a first stage of the second compressor and has a radius R 2 defined at a splitter rim, each radius R 1 , R 2 defined relative to the longitudinal axis, the splitter rim guides flow into the core flowpath; and
wherein a ratio of an axial length L 10 between the radii R 1 , R 2 to the radius R 2 is less than 1.2.
18. The method as recited in claim 17 , wherein the ratio of Rh to Rt is less than 0.30.
19. The method as recited in claim 17 , wherein: a ratio of R 1 to R 2 is greater than 0.5.
20. The method as recited in claim 19 , wherein the ratio of Rh to Rt is less than 0.30.
21. The method as recited in claim 20 , wherein the ratio of R 1 to R 2 is between 0.55 and 1.0.
22. The method as recited in claim 21 , comprising providing a geared architecture that rotates the fan section at a different speed than the fan drive turbine.
23. The method as recited in claim 17 , wherein the first turbine includes two stages.