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US 2020/0400071 A1Published application
Geared turbofan engine gearbox arrangement
AssigneeRaytheon Technologies Corporation
InventorsFrederick M. Schwarz, William G. Sheridan
Life of this patent
Estimating the term…
Abstract
A gas turbine engine according to the present disclosure includes, among other things, a fan section having a plurality of fan blades, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a core engine including a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section.
Claims20
A gas turbine engine comprising:
- a fan section having a plurality of fan blades, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh at a first inboard boundary of a first flowpath, and a ratio of Rh to Rt being less than 0.4;
- and
- a core engine including a first turbine configured to drive a first compressor, and a fan drive turbine configured to drive the fan section, the fan drive turbine including 5 stages, and the fan drive turbine including fewer stages than the first compressor.
A gas turbine engine comprising:
- a fan section having a plurality of fan blades configured to deliver airflow to a bypass duct, the plurality of fan blades having a peak tip radius Rt and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath, 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 configured to drive a first compressor, a second compressor, and a fan drive turbine configured to drive the fan section and arranged aft of the first turbine;
- and
- wherein a second inboard boundary of the core flowpath has a radius R1 defined at a first stage of the second compressor and a radius R2 defined at a splitter rim, each radius R1, R2 defined relative to the longitudinal axis, the splitter rim is configured to guide flow into the core flowpath, and a ratio of R1 to R2 is greater than 0.5.
1 more element
A method of designing a gas turbine engine comprising:
- providing a fan section having a plurality of fan blades, the fan blades having a peak tip radius Rt and an inboard leading edge radius Rh defined at a first inboard boundary of a first flowpath, and a ratio of Rh to Rt being less than 0.4;
- and
- providing a core engine including a first turbine configured to drive a first compressor, a second compressor, and a fan drive turbine configured to drive 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.
Dependencies are inferred from the claim text.