Geared turbofan architecture

1. A gas turbine engine comprising:

a fan section having a fan bypass ratio greater than 12 and less than 20;

a speed reduction device drivingly connected to the fan section;

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 , including an unshrouded turbine blade row adjacent the transition duct.

3. The gas turbine engine of claim 2 , wherein the unshrouded turbine blade row adjacent the transition duct is located in the high pressure turbine.

4. The gas turbine engine of claim 1 , wherein the fewer support struts include at least four equally spaced support struts.

5. The gas turbine engine of claim 1 , wherein the high pressure compressor includes at least 12 stages and a pressure ratio of at least 22:1 and less than 35:1.

6. The gas turbine engine of claim 1 , wherein the high pressure compressor includes at least 10 stages and a pressure ratio of at least 29:1 and less than 35:1.

7. The gas turbine engine of claim 1 , wherein the low pressure turbine is a drive turbine for the fan section.

8. The gas turbine engine of claim 1 , wherein the struts have an airfoil shape.

9. The gas turbine engine of claim 8 , wherein the struts are free of cooling fluid passages.

10. 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 fan section having a fan bypass ratio greater than 12 and less than 20 and a speed reduction device drivingly connected to the fan section.

11. The method as recited in claim 10 , 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.

12. The method as recited in claim 10 , including rotating a low pressure compressor three times as fast as a fan.

13. The method as recited in claim 10 , wherein the fewer support struts include at least four equally spaced support struts.

14. The method as recited in claim 10 , 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.

15. The gas turbine engine of claim 1 , 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.

16. The method of claim 10 , 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.

17. The gas turbine engine of claim 10 , 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.