Gas turbine engine compressor components comprising thermal barriers, thermal barrier systems, and methods of using the same

Abstract

Gas turbine engine compressor disks having a hot flowpath side; a shaft having a first surface positioned in the hot flowpath side; and a thermal barrier applied to at least the first surface of the shaft where the thermal barrier is operable to maintain the temperature of the shaft below about 700° C. (1300° F.) when the hot flowpath side experiences a service operating temperature of from about 700° C. (1300° F.) to about 788° C. (1450° F.).

1 . A gas turbine engine compressor disk comprising:
 a hot flowpath side;   a shaft having a first surface positioned in the hot flowpath side; and   a thermal barrier applied to at least the first surface of the shaft wherein the thermal barrier is operable to maintain the temperature of the shaft below about 700° C. (1300° F.) when the hot flowpath side experiences a service operating temperature of from about 700° C. (1300° F.) to about 788° C. (1450° F.).   
     
     
         2 . The disk of  claim 1  comprising a thermal barrier selected from the group consisting of thermal barrier coatings, metal heat shields, and thermal blankets. 
     
     
         3 . The disk of  claim 2  wherein the thermal barrier coating comprises yttria stabilized zirconia. 
     
     
         4 . The disk of  claim 2  wherein the metal heat shield comprises a material selected from the group consisting of metals, metal alloys, and metal superalloys based on nickel, cobalt, iron, and combinations thereof. 
     
     
         5 . The disk of  claim 2  wherein the thermal blanket comprises a low conductivity material. 
     
     
         6 . A thermal barrier system for a hot flowpath side of a gas turbine engine compressor shaft comprising:
 a thermal barrier selected from the group consisting of thermal barrier coatings, metal heat shields, and thermal blankets applied to at least a first surface of the compressor shaft wherein the thermal barrier is operable to maintain the temperature of the shaft below about 700° C. (1300° F.) when the hot flowpath side experiences a service operating temperature of from about 700° C. (1300° F.) to about 788° C. (1450° F.).   
     
     
         7 . The system of  claim 6  wherein the thermal barrier comprises a thermal expansion of greater than about 8×10 −6 /° C. 
     
     
         8 . The system of  claim 7  wherein the thermal barrier comprises an in-plane modulus of less than 5 Msi. 
     
     
         9 . The system of  claim 8  wherein the thermal barrier is capable of providing a reduction in the service operating temperature of the shaft of at least about 5° C./mm of thermal barrier thickness. 
     
     
         10 . The system of  claim 9  wherein the thermal barrier coating comprises yttria stabilized zirconia. 
     
     
         11 . The system of  claim 9  wherein the metal heat shield comprises a material selected from the group consisting of metals, metal alloys, and metal superalloys based on nickel, cobalt, iron, and combinations thereof. 
     
     
         12 . The system of  claim 9  wherein the thermal blanket comprises a low conductivity material. 
     
     
         13 . A method for reducing service temperature of operation of a gas turbine engine compressor shaft comprising:
 providing a compressor shaft having a first surface in a hot flowpath side and a second surface in a cool cavity side; and   applying a thermal barrier to at least the first surface of the compressor shaft wherein the thermal barrier is operable to maintain the temperature of the shaft below about 700° C. (1300° F.) when the hot flowpath side experiences a service operating temperature of from about 700° C. (1300° F.) to about 788° C. (1450° F.).   
     
     
         14 . The method of  claim 13  comprising providing a compressor shaft comprising a polycrystalline substrate selected from the group consisting of metals, metal alloys, and metal superalloys based on nickel, cobalt, iron, and combinations thereof. 
     
     
         15 . The method of  claim 14  comprising applying a thermal barrier capable of providing a reduction in the service operating temperature of the shaft of at least about 5° C./mm of thermal barrier thickness. 
     
     
         16 . The method of  claim 14  comprising applying a thermal barrier having a thermal expansion of greater than about 8×10 −6 /° C. 
     
     
         17 . The method of  claim 14  comprising applying a thermal barrier having an in-plane modulus of less than 5 Msi. 
     
     
         18 . The method of  claim 17  wherein the thermal barrier is selected from the group consisting of thermal barrier coatings, metal heat shields, and thermal blankets. 
     
     
         19 . The method of  claim 18  wherein the thermal barrier coating comprises yttria stabilized zirconia. 
     
     
         20 . The method of  claim 18  wherein the metal heat shield comprises a material selected from the group consisting of metals, metal alloys, and metal superalloys based on nickel, cobalt, iron, and combinations thereof.