Rotor for a turbomachine

Abstract

The invention relates to a rotor ( 10 ) for a turbomachine, in particular for a gas turbine, having a basic rotor body ( 12 ) and a plurality of blades ( 14 ), wherein at least one damping element ( 24 ) is provided in the circumferential direction between adjacent blades ( 14 ) of the rotor ( 10 ).

1 . A rotor ( 10 ) for a turbomachine, in particular for a gas turbine,
 having a basic rotor body ( 12 ) and a plurality of blades ( 14 ),   is hereby characterized in that   at least one damping element ( 24 ) is provided in the circumferential direction between adjacent blades ( 14 ) of the rotor ( 10 ).   
     
     
         2 . The rotor according to  claim 1 , further characterized in that the rotor ( 10 ) is an integrally bladed rotor ( 10 ), in particular a rotor ( 10 ) in which the basic rotor body ( 12 ) and blades ( 14 ) are welded, particularly friction-welded, directly to one another or via a separate intermediate piece. 
     
     
         3 . The rotor according to  claim 1 , further characterized in that the blades ( 14 ) have a blade neck ( 18 ), by means of which they are joined to the basic rotor body ( 12 ). 
     
     
         4 . The rotor according to  claim 3 , further characterized in that the blade necks ( 18 ) have extensions ( 22 ) that together form a shroud that bounds by its radially outward-pointing surface an annular flow channel of the turbomachine. 
     
     
         5 . The rotor according to  claim 3 , further characterized in that the damping element ( 24 ) is essentially disposed in the region of adjacent blade necks ( 18 ). 
     
     
         6 . The rotor according to one of  claim 3 , further characterized in that the blade necks ( 18 ) of adjacent blades ( 14 ) are distanced from one another in the circumferential direction and a free space is formed between the blade necks ( 18 ). 
     
     
         7 . The rotor according to  claim 6 , further characterized in that the damping element ( 24 ) forms a seal of the free space between adjacent blade necks ( 18 ) relative to a gas flow channel of the turbomachine. 
     
     
         8 . The rotor according to  claim 7 , further characterized in that the damping element ( 24 ) forms a part of the shroud and bounds the annular gas flow channel by its surface pointing radially outward. 
     
     
         9 . The rotor according to  claim 1 , further characterized in that the damping element ( 24 ) is disposed as an inserted piece between the blades ( 14 ). 
     
     
         10 . The rotor according to  claim 1 , further characterized in that the damping element ( 24 ) is joined to the adjacent blades ( 14 ) in a form-fitting manner, via a press connection and/or cohesively. 
     
     
         11 . The rotor according to  claim 1 , further characterized in that the damping element ( 24 ) is joined to the basic rotor body ( 12 ) in a form-fitting manner, via a press connection and/or cohesively. 
     
     
         12 . The rotor according to  claim 1 , further characterized in that a channel ( 26 ) is provided between the at least one damping element ( 24 ) and the basic rotor body ( 12 ). 
     
     
         13 . The rotor according to  claim 12 , further characterized in that the channel ( 26 ) is formed at least partially by a notch ( 28 ) in the basic rotor body ( 12 ) and/or in the blade neck ( 18 ). 
     
     
         14 . The rotor according to  claim 12 , further characterized in that the channel ( 26 ) is a cooling channel. 
     
     
         15 . The rotor according to  claim 1 , further characterized in that a plurality of damping elements ( 24 ) that are joined together are provided. 
     
     
         16 . The rotor according to  claim 1 , further characterized in that a plurality of damping elements ( 24 ) that are introduced on an axially attachable support ring is provided.