Transition duct for a gas turbine engine

Abstract

A transition duct for a gas turbine engine typically contains one or more non-turning fairings. These fairings disturb the air flow through the duct, leading to flow separation and aerodynamic inefficiency. The invention proposes non-axisymmetric perturbations in the duct end walls to minimise these undesirable aerodynamic effects. This permits the construction of shorter and lighter transition ducts, with more pronounced curvature.

1 . A transition duct for a gas turbine engine, the duct comprising circumferentially-extending radially inner and outer end walls defining A an annular flow path for a gas flow, the duct having different radii from the engine centreline at its upstream and downstream ends, the duct further comprising a radially extending fairing, in which at least one of the end walls has a profile defined by a plurality of non-axisymmetric perturbations at a succession of axial positions so as to reduce or prevent separation of the end wall boundary layer flow near to the fairing in use. 
     
     
         2 . A transition duct as in  claim 1 , in which the perturbations are applied to the radially inner end wall. 
     
     
         3 . A transition duct as in  claim 1  in which the fairing has a camber line and the perturbations are symmetrical about the camber line. 
     
     
         4 . A transition duct as in  claim 3 , in which the fairing is symmetrical about the camber line. 
     
     
         5 . A transition duct as in  claim 1 , in which at least some of the perturbations alter the local radius of curvature of the end wall. 
     
     
         6 . A transition duct as in  claim 5 , in which the local convex radius of curvature of the end wall, viewed in the circumferential direction, is increased in the region of the fairing leading edge and the local concave radius of curvature of the end wall, viewed in the circumferential direction, is increased in the region of the fairing trailing edge. 
     
     
         7 . A transition duct as in  claim 5 , in which the circumferential extent of the perturbations is between 5% and 100% of the fairing pitch. 
     
     
         8 . A transition duct as in  claim 5 , in which the circumferential extent of the perturbations is substantially 50% of fairing pitch. 
     
     
         9 . A transition duct as in  claim 5 , in which the maximum amplitude of the perturbations is between 2% and 15% of chord. 
     
     
         10 . A transition duct as in  claim 5 , in which the maximum amplitude of the perturbations is substantially 7% of chord. 
     
     
         11 . A transition duct as in  claim 1 , in which at least some of the perturbations locally alter the cross-sectional area of the duct close to the fairing surface. 
     
     
         12 . A transition duct as in  claim 11 , in which the cross-sectional area of the duct is reduced in the region of the fairing leading edge and increased in the region of the fairing trailing edge. 
     
     
         13 . A transition duct as in  claim 11 , in which the circumferential extent of the perturbations is between 5% and 100% of the fairing pitch. 
     
     
         14 . A transition duct as in  claim 11 , in which the circumferential extent of the perturbations is substantially 50% of fairing pitch. 
     
     
         15 . A transition duct as in  claim 11 , in which the maximum amplitude of the perturbations is between 2% and 15% of chord. 
     
     
         16 . A transition duct as in  claim 11 , in which the maximum amplitude of the perturbations is substantially 7% of chord. 
     
     
         17 . A transition duct as in  claim 1 , in which at least one spline fit is applied through circumferentially corresponding points at different axial locations to blend the perturbations smoothly into the axisymmetric end wall profile. 
     
     
         18 . A gas turbine engine incorporating a transition duct according to  claim 1 . 
     
     
         19 .- 20 . (canceled)