Sealing system for a turbomachine

Abstract

A sealing system for a turbomachine, in particular, for a gas turbine, having an annular space between a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades. The sealing system comprises at least one first sealing fin disposed on an end of the rotor blade facing the wall, at least one second abradable lining disposed following the first sealing fin in the flow direction on the end of the rotor blade facing the wall, at least one first abradable lining disposed on the inside of the wall and opposite the first sealing fin, and at least one second sealing fin disposed in the flow direction following the first abradable lining on the inside of the wall and opposite the second abradable lining. The invention furthermore relates to a gas turbine comprising at least one sealing system.

What is claimed is: 
     
         1 . A sealing system for a turbomachine having a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades, the flow-restricting wall and the at least one series of rotor blades defining an annular space therebetween having a flow direction, the sealing system comprising:
 a first seal and a second seal disposed in the annular space between the flow-restricting wall and the at least one series of rotor blades;   wherein the first seal includes
 a first sealing fin that is disposed on an end of a rotor blade facing the flow-restricting wall, the rotor blade being from the at least one series of rotor blades, and 
 a first abradable lining that is disposed on an inside of the flow-restricting wall and is opposite the first sealing fin; and 
   wherein the second seal is disposed following the first seal in the flow direction and includes
 a second abradable lining that is disposed on the end of the rotor blade facing the flow-restricting wall, and 
 a second sealing fin that is disposed on the inside of the flow-restricting wall and is opposite the second abradable lining. 
   
     
     
         2 . A sealing system in accordance with  claim 1 , wherein the first sealing fin and the second abradable lining are disposed on a shroud of the rotor blade. 
     
     
         3 . A sealing system in accordance with  claim 2 , wherein the shroud has at least one chamfer in a rearward region in the flow direction. 
     
     
         4 . A sealing system in accordance with  claim 3 , wherein the chamfer is at an angle a in the range between 10° and 65° relative to a longitudinal axis of the turbomachine. 
     
     
         5 . A sealing system in accordance with  claim 4 , wherein the chamfer is at an angle a in the range between 20° and 45° relative to a longitudinal axis of the turbomachine. 
     
     
         6 . Sealing system according to  claim 2 , characterized in that the first sealing fin is integrated into the shroud. 
     
     
         7 . A sealing system in accordance with  claim 1 , wherein the second abradable lining has at least one chamfer in a rearward region in the flow direction. 
     
     
         8 . A sealing system in accordance with  claim 7 , wherein the chamfer is at an angle a in the range between 10° and 65° relative to a longitudinal axis of the turbomachine. 
     
     
         9 . A sealing system in accordance with  claim 8 , wherein the chamfer is at an angle a in the range between 20° and 45° relative to a longitudinal axis of the turbomachine. 
     
     
         10 . A sealing system in accordance with  claim 1 , wherein at least one of the first sealing fin or the second sealing fin is tilted opposite the flow direction. 
     
     
         11 . A sealing system in accordance with  claim 1 , wherein the second sealing fin is integrated into the flow-restricting wall. 
     
     
         12 . A sealing system in accordance with  claim 1 , wherein the sealing system is provided in the form of a stepped labyrinth. 
     
     
         13 . A gas turbine aircraft engine comprising:
 a casing wall;   at least one series of rotor blades including a plurality of rotor blades, the casing wall and the at least one series of rotor blades defining an annular space therebetween having a flow direction;   a first seal disposed in the annular space between the casing wall and the at least one series of rotor blades;
 the first seal including a first sealing fin that is disposed on an end of a rotor blade facing the casing wall, the rotor blade being from the at least one series of rotor blades, and 
 a first abradable lining that is disposed on an inside of the casing wall and is opposite the first sealing fin; and 
   a second seal disposed in the annular space following the first seal in the flow direction and including
 a second abradable lining that is disposed on the end of the rotor blade facing the casing wall, and 
 a second sealing fin that is disposed on the inside of the casing wall and is opposite the second abradable lining. 
   
     
     
         14 . A gas turbine aircraft engine in accordance with  claim 13 , wherein the first sealing fin and the second abradable lining are disposed on a shroud of the rotor blade. 
     
     
         15 . A gas turbine aircraft engine in accordance with  claim 14 , wherein the first sealing fin is integrated into the shroud. 
     
     
         16 . A gas turbine aircraft engine in accordance with  claim 13 , wherein at least one of the first sealing fin or the second sealing fin is tilted opposite the flow direction. 
     
     
         17 . A gas turbine aircraft engine in accordance with  claim 13 , wherein the second sealing fin is integrated into the casing wall.