Blade cooling and sealing system

Abstract

A cooled component for a gas turbine engine, for example a turbine rotor blade, is provided. The component has an internal cooling flow passage. Cooling air is passed through the internal cooling flow passage to remove heat from the component and thereby reduce its temperature. The cooling air is bled from the internal cooling passage after it has passed through a portion of the passage into an internal bleed flow passage. This bled air is then used in a seal. Thus, some of the cooling air that enters the internal cooling flow passage is used both to cool the component and to form a seal, for example with another component.

1 . A blade or vane for a gas turbine engine comprising:
 a cooling fluid inlet configured to allow cooling fluid into the blade or vane;   a cooling fluid outlet configured to allow cooling fluid out of the blade or vane;   an aerofoil that, in use, is gas-washed by a working fluid of the gas turbine engine, the aerofoil extending from a platform;   an internal cooling flow passage between the cooling fluid inlet and the cooling fluid outlet configured to channel cooling fluid through the interior of the blade or vane;   an internal bleed flow passage in fluid communication with the internal cooling flow passage at a bleeding position, and configured to bleed at least a portion of the cooling fluid from the internal cooling flow passage for use as a sealing flow; and   a sealing flow outlet through which fluid from the internal bleed flow passage exits the blade or vane, the sealing flow outlet being positioned to allow the sealing flow to be used in a seal during operation, wherein:   the internal cooling flow passage extends through at least a part of the aerofoil before the sealing flow is bled off at the bleeding position;   the cooling fluid inlet is located radially inward of the platform; and   the sealing flow outlet is located radially inward of the platform.   
     
     
         2 . A blade or vane according to  claim 1 , wherein:
 the internal cooling flow passage is a multipass cooling flow passage, each pass being arranged to carry cooling fluid in either a radially outward direction or a radially inward direction; and   the bleeding position is after at least one pass through the internal cooling passage from the cooling fluid inlet.   
     
     
         3 . A blade or vane according to  claim 1 , wherein the internal cooling flow passage comprises an s-shape. 
     
     
         4 . A blade or vane according to  claim 1 , further comprising a fixture for attaching the blade or vane to a gas turbine engine, wherein:
 the cooling fluid inlet is located in the fixture; and   the sealing flow outlet is located in the fixture.   
     
     
         5 . A blade or vane according to  claim 1 , wherein the cooling flow outlet is located towards the radially outer end of the blade or vane. 
     
     
         6 . A blade or vane according to  claim 4 , wherein the bleeding position is located in the fixture. 
     
     
         7 . A blade or vane according to  claim 1 , wherein:
 the blade or vane, in use, has an upstream side and a downstream side defined relative to a flow direction through the gas turbine engine; and   the sealing flow outlet is located on the downstream side of the blade or vane.   
     
     
         8 . A blade or vane according to  claim 1 , wherein the bleeding position is at least 20% along the length of the internal cooling flow passage from the cooling fluid inlet. 
     
     
         9 . A blade or vane according to  claim 1 , wherein in the range of from 10% to 70% of the mass flow rate of cooling flow entering the blade through the cooling flow inlet is bled through the internal bleed flow passage and out through the sealing flow outlet. 
     
     
         10 . A blade or vane according to  claim 1 , wherein the internal bleed flow passage has a dog-leg shape such that it passes to the side of the internal cooling flow passage when viewed in the radial direction. 
     
     
         11 . A compressor or turbine for a gas turbine engine having at least one rotor stage having rotor blades and at least one stator stage having stator vanes, wherein at least one of the rotor blades or stator vanes is a blade or vane according to  claim 1 . 
     
     
         12 . A rotor-stator stage for a gas turbine, wherein: 
       the rotor stage comprises at least one blade or vane according to  claim 1 ; and
 the sealing flow exiting from the sealing flow outlet is used in a seal between the rotor stage and the stator stage. 
 
     
     
         13 . A method of using a flow in a gas turbine engine to cool a blade or a vane in a stage of the gas turbine engine and to seal between a rotor stage and a stator stage of the gas turbine engine, wherein:
 the blade or vane comprises an aerofoil that, in use, is gas-washed by a working fluid of the gas turbine engine; and   the method comprising using at least a part of the flow both to cool the aerofoil and to seal between the stages.   
     
     
         14 . A method according to  claim 13 , comprising:
 cooling the aerofoil by passing a cooling flow into the blade or vane, through an internal cooling passage inside the aerofoil, and out of the blade or vane through a cooling fluid outlet;   bleeding at least a part of the cooling flow passing through the internal cooling passage through a sealing flow outlet in the blade or vane before it reaches the cooling fluid outlet; and   using the flow passing out of the blade through the sealing flow outlet in a seal between the stages.   
     
     
         15 . A method according to  claim 13 , wherein the flow that is used for cooling and sealing is bled from a compressor of the gas turbine engine before entry into a combustor. 
     
     
         16 . A rotor-stator stage for a gas turbine, wherein: 
       the rotor stage comprises at least one blade and vane according to  claim 1 ; and
 the sealing flow exiting from the sealing flow outlet is used in a seal between the rotor stage and the stator stage.