Gas turbine engine with forward moment arm
AssigneeRaytheon Technologies Corporation
InventorsDavid Bomzer, Frederick M. Schwarz
Life of this patent
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Abstract
A gas turbine engine includes a plurality of fan blades rotatable about an axis, wherein each of the plurality of fan blades includes a leading edge. The gas turbine engine also includes turbine section includes an aft most turbine blade having a trailing edge and a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis. A center of gravity of the gas turbine engine is located a first axial distance from the trailing edge of the aft most turbine blade that is between about 35% and about 75% of a total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
Claims20
A gas turbine engine comprising:
- a plurality of fan blades rotatable about an axis, wherein each of the plurality of fan blades includes a leading edge, wherein the plurality of fan blades comprises no more than 20 fan blades;
- a compressor section at least partially defining a core flow path and a bypass flow path, wherein a bypass ratio is greater than ten;
- a turbine section including an aft most turbine blade, wherein the aft most turbine blade includes a trailing edge;
- and
- a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis;
- wherein a center of gravity of the gas turbine engine is located a first axial distance from the trailing edge of the aft most turbine blade that is between 40% and 70% of a length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
A turbofan engine comprising
- a plurality of fan blades rotatable about an axis, wherein each of the plurality of fan blades includes a leading edge;
- a bypass passage for a bypass airflow, wherein a bypass ratio is greater than ten;
- a turbine section including an aft most rotating turbine blade, wherein the aft most rotating turbine blade includes a trailing edge, wherein the turbine section includes a low pressure turbine section and a ratio between a number of the plurality of fan blades and a number of low pressure turbine rotors is between 3.3 and 8.6;
- and
- a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis;
- wherein an internal moment arm of the turbofan engine has a ratio of a first distance from a center of gravity of the turbofan engine to the trailing edge of the aft most rotating turbine blade to a length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade that is between 40% and 70%.
A method of assembling a gas turbine engine comprising:
- supporting a plurality of fan blades about an axis of rotation with each of the plurality of fan blades includes a leading edge;
- supporting a compressor section at least partially defining a core flow path and a bypass flow path, wherein a bypass ratio is greater than ten;
- supporting a turbine section including an aft most turbine blade, wherein each of the aft most turbine blade includes a trailing edge about the axis of rotation;
- supporting a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis;
- selecting components of the gas turbine engine structure to orientate a center of gravity of the gas turbine engine located a first axial distance from the trailing edge of the aft most turbine blade that is between 35% and 75% of a length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
Dependencies are inferred from the claim text.