Turbine vane assembly with cooling feature

Abstract

A turbine vane assembly adapted for use in a gas turbine engine includes a support and a turbine vane arranged around the support. The support is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support.

What is claimed is: 
     
         1 . A turbine vane assembly adapted for use in a gas turbine engine, the assembly comprising
 a vane made of ceramic matrix composite materials, the vane including an outer end wall, an inner end wall spaced radially inward of the outer end wall relative to a central reference axis to define a primary gas path therebetween, and an airfoil that extends from the outer end wall to the inner end wall across the primary gas path,   a spar made of metallic materials that is spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path, the spar including a mount panel engaged with the vane at at least one location radially spaced from the primary gas path to receive aerodynamic loads from the vane and a rod that extends radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path,   wherein the spar further includes a plurality of heat transfer augmentation features arranged at radial locations along the primary gas path that are configured to induce turbulence in cooling air supplied to the gap between the vane and the spar across the primary gas path during use of the turbine vane such that heat is more effectively transferred from the spar to the cooling air while avoiding conductive heat transfer from the ceramic matrix composite materials of the vane to the metallic materials of the spar that would be caused by contact between the vane and the spar across the primary gas path.   
     
     
         2 . The assembly of  claim 1 , wherein the plurality of heat transfer augmentation features include a plurality of protrusions that extend from the rod toward the airfoil of the vane. 
     
     
         3 . The assembly of  claim 2 , wherein the plurality of protrusions are located only along a side of the rod facing a pressure side of the airfoil. 
     
     
         4 . The assembly of  claim 2 , wherein the plurality of protrusions are located only along a side of the rod facing a suction side of the airfoil. 
     
     
         5 . The assembly of  claim 2 , wherein the plurality of heat transfer augmentation features further include a plurality of through holes that extend in a general circumferential direction, tangent to a circumference of the gas turbine engine with respect to the central reference axis through the rod in a portion of the rod which is a third of a chord length of the rod. 
     
     
         6 . The assembly of  claim 1 , wherein the plurality of heat transfer augmentation features include a plurality of depressions that extend inwardly into the rod of the spar. 
     
     
         7 . The assembly of  claim 6 , wherein the plurality of depressions are located only along a side of the rod facing a pressure side of the airfoil. 
     
     
         8 . The assembly of  claim 6 , wherein the plurality of depressions are located only along a side of the rod facing a suction side of the airfoil. 
     
     
         9 . The assembly of  claim 6 , wherein the plurality of heat transfer augmentation features further include a plurality of through holes that extend in a general circumferential direction, tangent to a circumference of the gas turbine engine with respect to the central reference axis through the rod in a portion of the rod which is a third of a chord length of the rod 
     
     
         10 . The assembly of  claim 1 , wherein the spar is formed to include a cooling air conduit that extends from outside the primary gas path into the rod which is formed to include cooling air holes sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path. 
     
     
         11 . The assembly of  claim 10 , wherein the cooling air holes are arranged to discharge cooling air toward a leading edge of the airfoil included in the vane to provide some cooling to the vane. 
     
     
         12 . A turbine vane assembly adapted for use in a gas turbine engine, the assembly comprising
 a vane made of ceramic matrix composite materials, the vane including an airfoil sized to extend in a radial direction relative to a central reference axis across a primary gas path of the gas turbine engine,   a spar made of metallic materials that is spaced from the airfoil of the vane at all radial locations across the primary gas path such that a gap is maintained between the vane and the spar across the primary gas path, the spar including (i) a mount panel engaged with the vane at a location radially spaced from the primary gas path, (ii) a rod that extends radially from the mount panel through a radially-extending passageway formed by the airfoil of the vane across the primary gas path, and (iii) a plurality of heat transfer augmentation features arranged at radial locations along the primary gas path configured to induce turbulence in cooling air supplied to the gap.   
     
     
         13 . The assembly of  claim 12 , wherein the plurality of heat transfer augmentation features include a plurality of protrusions that extend from the rod toward the airfoil of the vane. 
     
     
         14 . The assembly of  claim 13 , wherein the plurality of protrusions are located only along a side of the rod facing a pressure side of the airfoil. 
     
     
         15 . The assembly of  claim 13 , wherein the plurality of protrusions are located only along a side of the rod facing a suction side of the airfoil. 
     
     
         16 . The assembly of  claim 12 , wherein the plurality of heat transfer augmentation features include a plurality of depressions that extend inwardly into the rod of the spar. 
     
     
         17 . The assembly of  claim 16 , wherein the plurality of depressions are located only along a side of the rod facing a pressure side of the airfoil. 
     
     
         18 . The assembly of  claim 16 , wherein the plurality of depressions are located only along a side of the rod facing a suction side of the airfoil. 
     
     
         20 . The assembly of  claim 12 , wherein the spar is formed to include a cooling air conduit that extends from outside the primary gas path into the rod which is formed to include cooling air holes sized to discharge cooling air from the cooling air conduit into the gap between the vane and the spar along the primary gas path.