Granted Patent B2

US 8,066,479 · App. 12/753,934 · Granted Nov 29, 2011

Abstract

Claims (26)

1. A damper assembly of a gas turbine engine comprising:

a platform mountable to a disk through a support;

a first damper attached to said platform, said first damper engageable with a first blade attachable to the disk;

a second damper attached to said platform, said second damper engageable with a second blade attachable to the disk;

a first seal which extends toward the first blade; and

a second seal which extends toward the second blade, said first seal and said second seal radially outboard of said respective first damper and said second damper.

2. The assembly as recited in claim 1 , further comprising:

a first seal which extends toward the first blade and a second seal which extends toward the second blade, said first seal and said second seal radially outboard of said respective first damper and said second damper.

3. The assembly as recited in claim 1 , wherein said first damper and said second damper are flexed inboard toward the disk to provide a contact interface with the first blade and the second blade, respectively.

4. A section of a gas turbine engine comprising:

a disk;

a first blade attached to said disk;

a second blade attached to said disk;

a damper assembly which extends from said disk between said first blade and said second blade, said damper assembly includes a first damper which engages said first blade and a second damper which engages said second blade;

a first seal which extends toward the first blade; and

a second seal which extends toward the second blade, said first seal and said second seal radially outboard of said respective first damper and said second damper.

5. The section as recited in claim 4 , wherein said first seal, said second seal, said first damper and said second damper extend from a platform.

6. The section as recited in claim 5 , wherein said platform is mounted to said disk through a support.

7. The section as recited in claim 6 , wherein said support extends generally along an axis of rotation of said disk.

8. The section as recited in claim 4 , wherein said first damper and said second damper are flexed inboard toward said disk to provide a contact interface with said first blade and said second blade, respectively.

9. The section as recited in claim 8 , wherein said first damper and said second damper each define a blunt end in contact with said first blade and said second blade respectively.

10. The section as recited in claim 4 , wherein said first damper engages an intermediate portion of said first blade and said second damper engages an intermediate portion of said second blade, said intermediate portion of said first blade and said second blade generally separates a root portion and an airfoil portion to generally define an inner boundary of an air flow path.

11. The section as recited in claim 4 , wherein said first blade and said second blade are fan blades.

12. A method of damping a blade in a gas turbine engine comprising:

dissipating vibratory motion between a blade and a damper assembly in the form of friction heating; and

sealing an inner boundary of an air flow path between a first and a second blade radially outboard of the damper assembly using a first seal which extends toward the first blade and a second seal which extends toward the second blade, wherein the damper assembly includes a first damper engageable with the first blade and a second damper assembly engageable with the second blade.

Assignments (9)