Low weight large fan gas turbine engine

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan configured to deliver airflow to a bypass passage, and a core engine configured to rotate the fan. The core engine includes a high pressure turbine section configured to drive a high pressure compressor section, and a low pressure turbine section configured to drive the fan and a low pressure compressor section. The fan has a fan diameter, Dfan, and the low pressure turbine section has a turbine diameter, Dturb.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a fan configured to deliver airflow to a bypass passage;   a core engine configured to rotate the fan, the core engine including:
 a high pressure turbine section configured to drive a high pressure compressor section, and 
 a low pressure turbine section configured to drive the fan and a low pressure compressor section, the fan and the low pressure turbine section configured to rotate at a common speed and in a common direction; and 
   wherein the gas turbine engine has a bypass ratio of greater than about 10, the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.   
     
     
         2 . The gas turbine engine as recited in  claim 1 , wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section. 
     
     
         3 . The gas turbine engine as recited in  claim 1 , wherein the fan includes a plurality of fan blades, and the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades. 
     
     
         4 . The gas turbine engine as recited in  claim 1 , wherein the high pressure turbine section includes two stages. 
     
     
         5 . The gas turbine engine as recited in  claim 1 , wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section. 
     
     
         6 . The gas turbine engine as recited in  claim 1 , wherein the low pressure turbine section has a greater number of stages than the low pressure compressor section. 
     
     
         7 . The gas turbine engine as recited in  claim 6 , wherein the fan has a pressure ratio of less than about 1.45. 
     
     
         8 . The gas turbine engine as recited in  claim 7 , wherein the fan includes fewer than 20 fan blades. 
     
     
         9 . The gas turbine engine as recited in  claim 1 , wherein the low pressure turbine section has a pressure ratio of greater than about 5. 
     
     
         10 . A gas turbine engine comprising:
 a fan having fewer than 20 fan blades situated at an inlet of a bypass passage;   a core engine configured to rotate the fan, the core engine including:
 a high pressure turbine section configured to drive a high pressure compressor section, and 
 a low pressure turbine section configured to drive the fan and a low pressure compressor section, the low pressure turbine section having a greater number of stages than the low pressure compressor section; and 
   wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a turbine diameter, Dturb, and the fan diameter Dfan and the turbine diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65.   
     
     
         11 . The gas turbine engine as recited in  claim 10 , wherein the turbine diameter Dturb is defined by an outer case surface of the low pressure turbine section. 
     
     
         12 . The gas turbine engine as recited in  claim 10 , wherein the fan diameter Dfan is defined by outer peripheral surfaces of the fan blades. 
     
     
         13 . The gas turbine engine as recited in  claim 10 , wherein the high pressure turbine section includes two stages. 
     
     
         14 . The gas turbine engine as recited in  claim 10 , wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section. 
     
     
         15 . The gas turbine engine as recited in  claim 14 , wherein the fan has a pressure ratio of less than about 1.5. 
     
     
         16 . The gas turbine engine according to  claim 10 , wherein the gas turbine engine has a bypass ratio of greater than about 10. 
     
     
         17 . The gas turbine engine according to  claim 16 , wherein the low pressure turbine has a pressure ratio of greater than about 5. 
     
     
         18 . A method of designing a gas turbine engine comprising:
 providing a fan configured to deliver airflow to a bypass duct;   providing a core engine configured to rotate the fan, the core engine including:
 a high pressure turbine section configured to drive a high pressure compressor section, and 
 a low pressure turbine section configured to rotate the fan at a common speed and in a common direction; 
   wherein the fan has a fan diameter, Dfan, the low pressure turbine section has a diameter, Dturb, and the fan diameter Dfan, and the low turbine section diameter Dturb have an interdependence represented by a scalable ratio Dturb/Dfan that is between 0.5 and 0.65; and   wherein the fan is configured to deliver a portion of air into the core engine, and a portion of air into the bypass duct, and a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the core engine, is greater than 10.   
     
     
         19 . The method as recited in  claim 18 , wherein the fan has a pressure ratio of less than about 1.5 and the low pressure turbine has a pressure ratio of greater than about 5. 
     
     
         20 . The method as recited in  claim 18 , comprising a low pressure compressor section driven by the low pressure turbine section, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and includes fewer stages than the high pressure compressor section, and the high pressure turbine section includes two stages.