Bleed flow passage

Abstract

A liner wall insert is provided for a compressor rotor stage of a gas turbine engine. Several liner wall inserts are provided radially outboard of the tips of the rotor blades. The liner wall inserts have bleed flow channels formed therein. The bleed flow channels are arranged to remove flow from a trailing edge region of the stage and re-inject the bleed flow at an upstream region. The re-injected bleed flow alters the flow field around the tips of the rotor blades, for example the tip leakage flow. Thus, the bleed flow is used to improve the efficiency of the compressor rotor stage, and thus of the gas turbine engine.

1 . A liner wall insert for a compressor rotor stage of a gas turbine engine, the compressor rotor stage having rotor blades rotatable about an engine axis, and the liner wall insert comprising:
 a gas washed surface that forms a part of an outer flow boundary inside which the rotor blades rotate during use; and   a bleed flow passage extending from a bleed flow inlet to a bleed flow outlet, the bleed flow inlet and the bleed flow outlet both being formed in the gas washed surface, wherein:   the bleed flow inlet and bleed flow outlet are arranged such that, when the liner wall insert is assembled in the compressor rotor stage, the bleed flow inlet is axially downstream of the bleed flow outlet.   
     
     
         2 . A liner wall insert according to  claim 1 , wherein the cross sectional area of the bleed flow passage is smaller at the bleed flow outlet than at the bleed flow inlet. 
     
     
         3 . A liner wall insert according to  claim 1 , wherein the gas washed surface of the liner wall insert is a segment of a cylinder or a segment of a frusto-conical shape. 
     
     
         4 . A liner wall insert according to  claim 1 , wherein the bleed flow passage follows a path that is substantially parallel to the gas washed surface. 
     
     
         5 . A method of manufacturing a liner wall insert according to  claim 1 , comprising:
 manufacturing a radially inner portion of the liner wall insert;   manufacturing a radially outer portion of the liner wall insert; and   joining the radially inner portion and the radially outer portion together, wherein:   the radially inner portion comprises first surfaces that form part of the bleed flow passage;   the radially outer portion comprises second surfaces that form part of the bleed flow passage; and   the first surfaces and the second surfaces come together to form the bleed flow passage when the radially inner portion and radially outer portion are joined together.   
     
     
         6 . A method of manufacturing a liner wall insert according to  claim 1 , wherein:
 the method comprises metal injection moulding to manufacture at least a part of the insert.   
     
     
         7 . A method of manufacturing a liner wall insert according to  claim 5 , wherein:
 the method comprises metal injection moulding to manufacture at least a part of the insert.   
     
     
         8 . A compressor rotor stage comprising:
 a plurality of rotor blades rotatable about an engine axis, the rotor blades extending to a radially outer tip; and   a plurality of liner wall inserts according to  claim 1 , wherein   the plurality of liner wall inserts are attached together so as to define an outer flow boundary inside which the tips of the rotor blades pass during use.   
     
     
         9 . A compressor rotor stage according to  claim 8 , wherein:
 the axial location of the bleed flow inlet of each liner wall insert corresponds to a trailing edge region of the rotor blades; and   the axial location of the bleed flow outlet of each liner wall insert corresponds to a leading edge region of the rotor blades.   
     
     
         10 . A compressor rotor stage according to  claim 8 , wherein the axial location of each bleed flow outlet is upstream of the axial location of the leading edge of the rotor blades. 
     
     
         11 . A compressor rotor stage according to  claim 8 , wherein the plurality of liner wall inserts are brazed together to form a ring. 
     
     
         12 . A compressor rotor stage according to  claim 8 , further comprising a casing radially outboard of the liner wall inserts, wherein:
 the casing and the liner wall inserts have cooperating location features that are engaged so as to hold the liner wall inserts in position.   
     
     
         13 . A compressor rotor stage according to  claim 8 , further comprising a casing radially outboard of the liner wall inserts, wherein:
 the liner wall inserts are joined to the casing so as to be held in position.   
     
     
         14 . A gas turbine engine comprising at least one compressor rotor stage according to  claim 8 . 
     
     
         15 . A method of improving the operation of a compressor rotor stage of a gas turbine engine, the compressor rotor stage comprising rotor blades ( 210 ) extending from a root to a tip and being rotatable about an engine axis within an outer flow boundary, the method comprising:
 bleeding flow from the compressor flow stream into a bleed flow inlet, through a bleed flow passage, and back out into the flow stream from a bleed flow outlet that is axially upstream of the bleed flow inlet, wherein:   the outer flow boundary is formed by a plurality of liner wall inserts joined together to circumferentially surround the rotor blades; and   both the bleed flow inlet and the bleed flow outlet are formed in the outer flow boundary.   
     
     
         16 . A method according to  claim 15 , wherein a plurality of the liner wall inserts have a bleed flow passage formed therein.