Systems and methods involving localized stiffening of blades

1. A gas turbine engine blade comprising:

a first recess located in a pressure side surface of the blade;

a second recess located in a suction side surface of the blade;

a first material positioned at least partially within the first recess such that the first material provides a first localized increase in stiffness of the blade;

a second material positioned at least partially within the second recess such that the second material provides a second localized increase in stiffness of the blade;

wherein the first recess and the second recess are oriented substantially not parallel to each other.

2. The blade of claim 1 , wherein the first material is a composite material comprising fibers.

3. The blade of claim 2 , wherein:

the first recess exhibits a major axis; and

the fibers are substantially aligned with the major axis of the first recess.

4. The blade of claim 1 , wherein:

the first material is mounted flush with the pressure side surface of the blade.

5. The blade of claim 1 , wherein the blade is formed of titanium and one of the first material or the second material comprises a titanium metal matrix composite.

6. The blade of claim 1 , wherein the blade is a fan blade.

7. The blade of claim 1 , wherein the second material is a composite material comprising fibers.

8. The blade of claim 7 , wherein:

the second recess exhibits a major axis; and

the fibers are substantially aligned with the major axis of the second recess.

9. The blade of claim 1 , wherein:

the second material is mounted flush with the suction side surface of the blade.

10. A gas turbine engine comprising:

a blade having a pressure side surface and a suction side surface;

a first recess located in the pressure side surface of the blade;

a second recess located in the suction side surface of the blade;

a first material positioned at least partially within the first recess such that the material provides a first localized increase in stiffness of the blade;

a second material position at least partially within the second recess such that the material provides a second localized increase in stiffness of the blade; and

wherein the first recess and the second recess are oriented substantially not parallel to each other.

11. The engine of claim 10 , wherein:

the engine comprises a fan; and

the blade is a blade of the fan.

12. The engine of claim 10 , further comprising a differential gear operative to drive the fan.

13. A method comprising:

stiffening discrete portions of a blade of a gas turbine engine such that aeroelastic tuning of the blade is facilitated, wherein stiffening comprises:

forming a first recess in a pressure side surface of the blade and a second recess in a suction side surface of the blade, wherein the first recess and the second recess are oriented substantially not parallel to each other;

positioning a first material in the first recess to selectively stiffen the blade in a vicinity of the first recess;

positioning a second material in the second recess to selectively stiffen the blade in a vicinity of the second recess.

14. The method of claim 13 , wherein forming the first recess comprises:

providing the blade without the first recess; and

producing the first recess in the pressure side surface of the blade.

15. The method of claim 13 , wherein one of the first material or the second material is a composite material comprising fibers.

16. The method of claim 15 , wherein the composite material is a silicon carbide fiber tape.

17. The method of claim 13 , wherein, in stiffening the discrete portions of a blade, a tendency of the blade to exhibit flutter during use is reduced.

18. The method of claim 13 , wherein forming the second recess comprises:

providing the blade without the second recess; and

producing the second recess in the suction side surface of the blade.