Blade platform with damper restraint
AssigneeRaytheon Technologies Corporation
InventorsCharles Thistle, Mohamed Hassan
Life of this patent
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Abstract
A gas turbine engine, having: a disk; a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with a platform of an adjacent blade that is secured to the disk; a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades; and a damper restraint located on an interior surface of each platform, wherein the damper restraint extends into the cavity and is a raised feature configured to contact a peripheral edge portion of a damper seal when it is adjacent to the gap defined by the platforms of the blades.
Claims16
A section of a gas turbine engine, comprising:
- a disk;
- a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with another platform of an adjacent blade that is secured to the disk;
- a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades;
- and
- a damper restraint located on an interior surface of each platform, wherein the damper restraint of each platform extends into the cavity and is a raised feature configured to contact a peripheral edge portion of one of a pair of tab portions extending from opposite sides of the damper seal and a peripheral edge of the damper seal adjacent to the pair of tab portions when the damper seal is adjacent to the gap defined by the platforms of the blades, and wherein the damper restraint limits radial, tangential and rotational movement of the damper seal when the damper seal is adjacent to the gap defined by the platforms of the blades.
A gas turbine engine, comprising:
- a compressor section;
- a turbine section;
- a disk, wherein the disk is a portion of the compressor section or a portion of the turbine section;
- a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with another platform of an adjacent blade that is secured to the disk;
- a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades;
- and
- a damper restraint located on an interior surface of each platform, wherein the damper restraint of each platform extends into the cavity and is a raised feature configured to contact a peripheral edge portion of one of a pair of tab portions extending from opposite sides of the damper seal and a peripheral edge of the damper seal adjacent to the pair of tab portions when the damper seal is adjacent to the gap defined by the platforms of the blades, and wherein the damper restraint limits radial, tangential and rotational movement of the damper seal when the damper seal is adjacent to the gap defined by the platforms of the blades.
Dependencies are inferred from the claim text.