Geared turbofan architecture

Abstract

A gas turbine engine includes a propulsor. A speed reduction device is drivingly connected to the propulsor. A compressor section with a high pressure compressor has between 8 and 13 stages and a pressure ratio of at least 16:1 and less than 35:1. A turbine section including a transition duct is located between a high pressure turbine and a low pressure turbine including fewer support struts than vanes in a first vane row of the low pressure turbine. The first vane row of the low pressure turbine is located downstream of the transition duct and downstream of a first row of blades in the low pressure turbine. The first row of blades in the low pressure turbine are immediately downstream of the support struts.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a propulsor;   a speed reduction device drivingly connected to the propulsor;   a compressor section with a high pressure compressor having between 8 and 13 stages and a pressure ratio of at least 16:1 and less than 35:1; and   a turbine section including a transition duct located between a high pressure turbine and a low pressure turbine including fewer support struts than vanes in a first vane row of the low pressure turbine;   wherein the first vane row of the low pressure turbine is located downstream of the transition duct and downstream of a first row of blades in the low pressure turbine;   wherein the first row of blades in the low pressure turbine are immediately downstream of the support struts.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein the low pressure turbine includes at least three stages. 
     
     
         3 . The gas turbine engine of  claim 2 , wherein the low pressure turbine is a three stage low pressure turbine. 
     
     
         4 . The gas turbine engine of  claim 2 , wherein the low pressure turbine includes no more than five stages. 
     
     
         5 . The gas turbine engine of  claim 2 , wherein the high pressure turbine is a two-stage high pressure turbine. 
     
     
         6 . The gas turbine engine of  claim 2 , including an unshrouded turbine blade row adjacent the transition duct. 
     
     
         7 . The gas turbine engine of  claim 6 , wherein the unshrouded turbine blade row adjacent the transition duct is located in the high pressure turbine. 
     
     
         8 . The gas turbine engine of  claim 2 , wherein the fewer support struts include at least four equally spaced support struts. 
     
     
         9 . The gas turbine engine of  claim 2 , wherein the high pressure compressor includes at least 8 stages, is located immediately upstream of a combustor, and includes the pressure ratio of at least 16:1 and less than 35:1. 
     
     
         10 . The gas turbine engine of  claim 2 , wherein the high pressure compressor includes at least 10 stages and a pressure ratio of at least 29:1 and less than 35:1. 
     
     
         11 . The gas turbine engine of  claim 2 , wherein the low pressure turbine is a drive turbine for the propulsor. 
     
     
         12 . The gas turbine engine of  claim 2 , wherein the propulsor is at least partially surrounded by an outer housing defining bypass ratio greater than 12 and less than 20. 
     
     
         13 . The gas turbine engine of  claim 2  wherein the struts have an airfoil shape and the struts are free of cooling fluid passages. 
     
     
         14 . The gas turbine engine of  claim 2 , wherein the vanes in the first vane row of the low pressure turbine are downstream of blades in the first blade row of the low pressure turbine. 
     
     
         15 . A method of designing a gas turbine engine for improved performance comprising:
 positioning a transition duct between a high pressure turbine and a low pressure turbine, wherein the transition duct includes fewer support struts than vanes in a first vane row of the low pressure turbine and the first vane row of the low pressure turbine is located downstream of the transition duct and a first row of blades in the low pressure turbine are immediately downstream of the support struts;   decreasing a design temperature in the transition duct by increasing a pressure ratio in a high pressure compressor in a compressor section and increasing extraction from a working fluid with the high pressure turbine; and   wherein the gas turbine engine includes a propulsor and a speed reduction device drivingly connected to the propulsor.   
     
     
         16 . The method of  claim 15 , wherein the low pressure turbine includes at least three stages and no more than five stages and the high pressure turbine is a two-stage high pressure turbine. 
     
     
         17 . The method as recited in  claim 16 , including decreasing a pressure ratio in a low pressure compressor when increasing the pressure ratio in the high pressure compressor to decrease the temperature in the transition duct. 
     
     
         18 . The method of  claim 16 , wherein the first vane row of the low pressure turbine is located downstream of the transition duct and downstream of the first row of blades in the low pressure turbine. 
     
     
         19 . The method as recited in  claim 16 , wherein the fewer support struts include at least four equally spaced support struts. 
     
     
         20 . The method as recited in  claim 16 , wherein the high pressure compressor includes at least 8 stages, is located immediately upstream of a combustor, and includes the pressure ratio of at least 16:1 and less than 35:1.