Rotor assembly for gas turbine engines

1. A gas turbine engine comprising:

a fan section including a fan shaft rotatable about an engine longitudinal axis;

at least one bearing assembly supporting the fan shaft; and

wherein the fan section includes a rotor assembly, the rotor assembly comprising:

a rotatable hub including a main body mechanically attached to the fan shaft, and including an array of annular flanges extending about an outer periphery of the main body to define an array of annular channels along the engine longitudinal axis, wherein each of the annular channels receives a composite reinforcement member that extends about the outer periphery;

an array of airfoils each including an airfoil section extending from a root section; and

a plurality of retention pins, each of the retention pins extending through the root section of a respective one of the airfoils, across the annular channels at a position radially outward of the respective composite reinforcement member, and through the annular flanges to mechanically attach each root section to the hub; and

a fan drive turbine that drives the fan shaft through a geared architecture, wherein the at least one bearing assembly is positioned radially outward of the outer periphery of the hub with respect to the engine longitudinal axis, and the at least one bearing assembly supports the fan shaft at a position radially outward of the geared architecture with respect to the engine longitudinal axis.

2. The gas turbine engine as recited in claim 1 , wherein the at least one bearing assembly includes a plurality of bearing assemblies distributed along the fan shaft, and a radially innermost one of the bearing assemblies supports the fan shaft at a position radially outward of the geared architecture with respect to the engine longitudinal axis.

3. The gas turbine engine as recited in claim 1 , wherein each airfoil section includes a metallic sheath and a composite core, each core including first and second ligaments at least partially received in respective internal channels defined in the respective sheath.

4. The gas turbine engine as recited in claim 3 , wherein, for each of the airfoils:

each one of the ligaments includes at least one interface portion in the root section that receives a respective one of the retention pins, and each one of the ligaments includes at least one composite layer that loops around the at least one interface portion such that opposed end portions of the at least one composite layer are joined together along the airfoil portion.

5. The gas turbine engine as recited in claim 3 , wherein each of the retention pins includes a plurality of segments linked together by and slideably received on an elongated carrier, the plurality of segments includes a first set of segments and a second set of segments distributed along a pin axis of the carrier, and each segment of the first set of segments is received in a respective one of the annular flanges and each segment of the second set of segments is received in a respective one of the first and second ligaments in an installed position.

6. The gas turbine engine as recited in claim 5 , wherein the pin axis is curved in the installed position.

7. The gas turbine engine as recited in claim 1 , wherein each of the annular flanges includes an array of retention members distributed circumferentially about the engine longitudinal axis, and each of the retention members includes a retention body that extends outwardly from a respective one of the annular flanges and has a contact surface dimensioned to abut against a sidewall of the airfoil section of a respective one of the airfoils to oppose circumferential movement of the respective airfoil.

8. The gas turbine engine as recited in claim 7 , wherein each of the retention members has an L-shaped geometry including a terminal end defining the contact surface.

9. The gas turbine engine as recited in claim 1 , wherein each composite reinforcement member is a carbon tape that is wound around the outer periphery two or more times in the respective one of the annular channels.

10. The gas turbine engine as recited in claim 9 , wherein:

the at least one bearing assembly includes a plurality of bearing assemblies distributed along the fan shaft, and a radially innermost one of the bearing assemblies supports the fan shaft at a position that is radially outward of the geared architecture with respect to the engine longitudinal axis; and

at least a portion of one of the bearing assemblies is positioned radially outward of the outer periphery of the hub with respect to the engine longitudinal axis.

11. The gas turbine engine as recited in claim 10 , wherein each composite reinforcement member defines a first thickness, and the hub defines a second thickness along the outer periphery that defines a respective one of the annular channels, and the second thickness is less than the first thickness.

12. The gas turbine engine as recited in claim 1 , wherein the fan section delivers a portion of airflow into a compressor section and another portion of airflow into a bypass duct.