Balanced composite root region for a blade of a gas turbine engine

1. A blade for a gas turbine engine, comprising:

an inner ply layer group comprising a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and

a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter, wherein the platform shell comprises a platform shell center of gravity located outboard of a center point of a contact width at which the platform shell root region contacts a disk attachment lug of an associated blade slot, the platform shell center of gravity is located to create a net moment operable to rotate the platform shell platform region toward the inner ply layer group, wherein the platform shell neck region at the platform shell platform region roll inward toward the inner ply layer group during operation of the gas turbine engine such that the platform shell neck region operates in compression.

2. The blade as recited in claim 1 , wherein the platform shell comprises a multiple of pieces that are assembled together around the inner ply layer group.

3. A blade for a gas turbine engine, comprising:

an inner ply layer group comprising a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and

a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter, wherein the platform shell neck region flares outwardly to intersect with the platform shell platform region at the platform shell neck region perimeter, the platform shell platform region comprises a first thickness inboard of the platform shell neck region perimeter and a second thickness outboard of the platform shell neck region perimeter, the first thickness different than the second thickness.

4. A blade for a gas turbine engine, comprising:

an inner ply layer group comprising a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and

a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter, wherein the platform shell neck region flares outwardly to intersect with the platform shell platform region at the platform shell neck region perimeter, the platform shell platform region comprises a first thickness inboard of the platform shell neck region perimeter and a second thickness outboard of the platform shell neck region perimeter, wherein the first thickness inboard of the platform shell neck region perimeter defines a first center of gravity at a first cross-section, and the second thickness outboard of the platform shell neck region perimeter defines a second center of gravity at the first cross-section, the first center of gravity and the second center of gravity positioned to minimize a rolling moment located at a point Z at the first cross-section on the platform shell neck region perimeter during operation of the gas turbine engine.

5. A blade for a gas turbine engine, comprising:

an inner ply layer group comprising a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and

a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter; and

a low density ceramic foam between the platform shell neck region and the inner ply layer group.

6. A blade for a gas turbine engine, comprising:

an inner ply layer group comprising a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and

a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, wherein the platform shell neck region at the platform shell platform region roll inward toward the inner ply layer group during operation of the gas turbine engine such that the platform shell neck region operates in compression.

7. The blade as recited in claim 6 , wherein the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter.

8. The blade as recited in claim 7 , wherein the platform shell neck region perimeter surrounds an airfoil opening perimeter.

9. The blade as recited in claim 8 , wherein the platform shell comprises a platform shell center of gravity located outboard of a center point of a contact width at which the platform shell root region contacts a disk attachment lug of an associated blade slot.

10. The blade as recited in claim 8 , wherein the platform shell comprises a platform shell center of gravity located outboard of a center point of a contact width at which the platform shell root region contacts a disk attachment lug of an associated blade slot.

11. The blade as recited in claim 10 , wherein the platform shell center of gravity is located to create a net moment operable to rotate the platform shell platform region toward the inner ply layer group.

12. The blade as recited in claim 8 , wherein the platform shell neck region flares outwardly to intersect with the platform shell platform region at the platform shell neck region perimeter, the platform shell platform region comprises a first thickness inboard of the platform shell neck region perimeter and a second thickness outboard of the platform shell neck region perimeter, the first thickness different than the second thickness.

13. The blade as recited in claim 8 , wherein the platform shell neck region flares outwardly to intersect with the platform shell platform region at the platform shell neck region perimeter, the platform shell platform region comprises a first thickness inboard of the platform shell neck region perimeter and a second thickness outboard of the platform shell neck region perimeter, wherein the first thickness inboard of the platform shell neck region perimeter defines a first center of gravity at a first cross-section, and the second thickness outboard of the platform shell neck region perimeter defines a second center of gravity at the first cross-section, the first center of gravity and the second center of gravity positioned to minimize a rolling moment located at a point Z at the first cross-section on the platform shell neck region perimeter.

14. A method of manufacturing a rotor blade of a gas turbine engine, comprising:

determining a platform shell center of gravity to be outboard of a center point of a contact width at which a platform shell root region contacts a disk attachment lug of an associated blade slot; and

adjusting an intersection of a platform shell neck region that flares outwardly away from a neck region of an inner ply region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening to minimize a local bending moment.

15. The method as recited in claim 14 , wherein the platform shell neck region is between a platform shell root region and the platform shell platform region.

16. The method as recited in claim 15 , wherein the platform shell neck region is between a platform shell root region and the platform shell platform region, the platform shell root region sheathing a flared region that defines the center point of the area of contact.

17. The method as recited in claim 14 , wherein adjusting the intersection further comprises adjusting the platform shell platform region to roll inward toward the inner ply layer group during operation of the gas turbine engine.

18. The method as recited in claim 17 , wherein the platform shell neck region operates in compression.