Extended root region and platform over-wrap for a blade of a gas turbine engine
A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.
1. A rotor blade for a gas turbine engine, comprising:
an inner ply layer group that at least partially defines a base of a root region, wherein the base is a bottommost surface of the root region;
a platform that comprises at least two platform sections, which, in combination, define the platform and surround the inner ply layer group; and
a platform over-wrap around the platform, the platform over-wrap formed as a continuous band that wraps around the at least two platform sections.
2. The blade as recited in claim 1 , wherein the inner ply layer group at least partially defines an airfoil.
3. The blade as recited in claim 1 , wherein the inner ply layer group, the platform, and the platform over-wrap comprise a flared surface of the root region of the rotor blade.
4. The blade as recited in claim 1 , wherein the inner ply layer group and the platform over-wrap are manufactured of a ceramic matrix material.
5. The blade as recited in claim 1 , wherein the inner ply layer group, the platform, and the platform over-wrap comprise a ceramic matrix material.
6. The blade as recited in claim 1 , wherein the platform over-wrap defines an attachment surface at which a disk interfaces.
7. The blade as recited in claim 6 , wherein the attachment surface defines a zone of compression through the platform over-wrap and into said inner ply layer group.
8. The blade as recited in claim 7 , wherein the zone of compression is defined along a line that is oriented at 30-50 degrees with respect to a blade root centerline.
9. The blade as recited in claim 7 , wherein a point on a centerline of the blade is defined by a line perpendicular to an attachment surface and extending from a point in the zone of compression of the attachment surface, the base of the root region defined below the point on the centerline.
10. The blade as recited in claim 9 , wherein the point in the zone of compression is a central point on the attachment surface that defines the zone of compression.
11. The blade as recited in claim 9 , wherein the point in the zone of compression is a bottommost point on the attachment surface that defines the zone of compression.
12. The blade as recited in claim 1 , wherein the base is perpendicular to a centerline of the blade.
13. The blade as recited in claim 1 , wherein an outer surface of the platform over-wrap below an attachment surface is parallel to a blade root centerline.
14. The blade as recited in claim 1 , wherein an outer surface of the platform over-wrap below an attachment surface forms an inward angle toward a centerline of up to about 10 degrees.
15. A method of manufacturing a rotor blade of a gas turbine engine, comprising:
assembling a platform around an inner ply layer group, the platform and the inner ply layer group at least partially forming a base of a root region, wherein the base is a bottommost surface of the root region;
wrapping the platform with a platform over-wrap, the platform comprises at least two platform sections which, in combination, define the platform and surround the inner ply layer group, the platform over-wrap at least partially forming the base; and
forming the platform over-wrap as a continuous band that wraps around the at least two platform sections.
16. The method as recited in claim 15 , further comprising forming an attachment surface with the platform over-wrap.
17. The method as recited in claim 15 , further comprising applying a silicon plasma spray to form a machinable layer on an attachment face.
18. The method as recited in claim 15 , further comprising forming a flared surface with the platform, the inner ply layer group, and the platform over-wrap.