Gas turbine engine combustor liner assembly with convergent hyperbolic profile

Abstract

A liner assembly for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a support shell with a convex profile which faces the heat shield. A further embodiment of the foregoing embodiment of the present disclosure is where the convex profile is defined by a hyperbolic cosine function. A further embodiment of any of the foregoing embodiments of the present disclosure is where the convex profile provides an approximate 4.5 inlet-to-exit area ratio. A further embodiment of any of the foregoing embodiments, of the present disclosure wherein the convex profile provides a flow acceleration toward approximately 0.5 Mach towards an end of a convergent section.

What is claimed is: 
     
         1 . A liner assembly for a combustor of a gas turbine engine comprising:
 a heat shield; and   a support shell with a convex profile which faces said heat shield.   
     
     
         2 . The liner assembly as recited in  claim 1 , wherein said convex profile is defined by a hyperbolic cosine function. 
     
     
         3 . The liner assembly as recited in  claim 1 , wherein said convex profile provides an approximate 4.5 inlet-to-exit area ratio. 
     
     
         4 . The liner assembly as recited in  claim 1 , wherein said convex profile provides a flow acceleration toward approximately 0.5 Mach towards a end of a convergent section. 
     
     
         5 . The liner assembly as recited in  claim 1 , further comprising an exit splitter that extends from said heat shield. 
     
     
         6 . The liner assembly as recited in  claim 5 , wherein said exit splitter is zigzag in shape. 
     
     
         7 . The liner assembly as recited in  claim 5 , further comprising a film hole located in a valley on each side of said exit splitter. 
     
     
         8 . The liner assembly as recited in  claim 1 , further comprising a plurality of studs which extend from said heat shield and are received through said support shell, said stud include a frustro-conical section. 
     
     
         9 . The liner assembly as recited in  claim 1 , wherein said heat shield includes a number of film holes which are approximately equal to a number of impingement holes through said support shell. 
     
     
         10 . The liner assembly as recited in  claim 1 , wherein said heat shield includes a multiple of pin fins. 
     
     
         11 . The liner assembly as recited in  claim 10 , wherein said multiple of pin fins are diamond-shaped. 
     
     
         12 . The liner assembly as recited in  claim 1 , wherein said heat shield includes a multiple of hemi-spherical dimples. 
     
     
         13 . The liner assembly as recited in  claim 12 , wherein said multiple of hemi-spherical dimples decrease in diameter toward an exit splitter. 
     
     
         14 . The liner assembly as recited in  claim 12 , wherein a center of said sphere of each of said multiple of hemi-spherical dimples are further displaced from an inner surface of said heat shield toward an exit splitter. 
     
     
         15 . A liner assembly for a combustor of a gas turbine engine comprising:
 a heat shield; and   a support shell non-parallel to said heat shield.   
     
     
         16 . The liner assembly as recited in  claim 15 , wherein said support shell defines a convex profile defined by a hyperbolic cosine function. 
     
     
         17 . A method of increasing pressure in a liner assembly of a combustor for a gas turbine engine, comprising:
 directing an airflow in a generally circumferential direction along a convergent flow channel within a cavity between a heat shield and a support shell.   
     
     
         18 . The method as recited in  claim 17 , further comprising:
 defining the convergent flow channel by a hyperbolic cosine function.   
     
     
         19 . The method as recited in  claim 17 , further comprising:
 defining the convergent flow channel to provide an approximate 4.5 inlet-to-exit area ratio.   
     
     
         20 . The method as recited in  claim 17 , further comprising:
 accelerating the airflow toward approximately 0.5 Mach towards an end of said convergent section.