Geared turbofan gas turbine engine architecture

Abstract

A gas turbine engine according to an example of the present disclosure includes a turbine section having a very high speed fan drive turbine such that a quantity defined by the exit area of the fan drive turbine multiplied by the square of the fan drive turbine rotational speed compared to the same parameters for a second turbine is at a ratio between 0.5 and 1.5. The engine includes a power density greater than 1.5 lbf/in 3 .

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a fan section including a fan rotor and fan, the fan including a plurality of fan blades, with a low fan pressure ratio in operation of less than 1.45, the low fan pressure ratio measured across the fan blades alone;   a compressor section;   wherein the fan delivers a portion of air into a bypass duct, and a bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section, and the bypass ratio is greater than 10;   a gear train having a gear ratio greater than 2.3;   a combustor in fluid communication with the compressor section;   a turbine section in fluid communication with the combustor and including a fan drive turbine and a second turbine, the fan drive turbine driving the fan rotor through the gear train wherein:
 the engine includes a power density greater than 1.5 lbf/in 3 , and power density is defined as Sea Level Takeoff Thrust in lbf produced divided by the volume of the turbine section in cubic inches; 
 the fan drive turbine includes at least one rotor having a bore radius (R), a live rim radius (r) and a bore width (W), and a ratio of r/R is between 2.00 and 2.30, and a ratio of r/W is between 4.65 and 5.55; 
 the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio greater than 5 in operation, wherein the fan drive turbine pressure ratio is a ratio of a pressure measured prior to the inlet as related to a pressure at the outlet prior to any exhaust nozzle; 
 the fan drive turbine has a first exit area and rotates at a first speed, the second turbine has a second exit area and rotates at a second speed, which is faster than the first speed, wherein a first performance quantity is defined as the product of the first speed squared and the first exit area, a second performance quantity is defined as the product of the second speed squared and the second exit area, and a performance ratio of the first performance quantity to the second performance quantity is between 0.5 and 1.5; and 
 the fan drive turbine has a number of fan drive turbine stages, and a ratio between the number of fan blades and the number of fan drive turbine stages is between 2.5 and 8.5. 
   
     
     
         2 . The gas turbine engine set forth in  claim 1 , wherein the power density is greater than or equal to 3.0 lbf/in 3 . 
     
     
         3 . The gas turbine engine set forth in  claim 2 , wherein the power density is greater than or equal to 4.0 lbf/in 3 . 
     
     
         4 . The gas turbine engine set forth in  claim 1 , wherein the second turbine is a two stage turbine. 
     
     
         5 . The gas turbine engine set forth in  claim 4 , wherein the performance ratio is greater than or equal to 0.8. 
     
     
         6 . The gas turbine engine set forth in  claim 5 , wherein the performance ratio is greater than or equal to 1.0. 
     
     
         7 . The gas turbine engine set forth in  claim 5 , wherein the power density is greater than or equal to 4.0 lbf/in 3 . 
     
     
         8 . The gas turbine engine set forth in  claim 5 , wherein the fan section has a low corrected fan tip speed less than 1150 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed divided by [(Tram° R)/(518.7° R)] 0.5 . 
     
     
         9 . The gas turbine engine set forth in  claim 8 , wherein the gear train is a planetary gear system and the power density is less than or equal to 5.5 lbf/in 3 . 
     
     
         10 . The gas turbine engine set forth in  claim 4 , wherein the plurality of fan blades is less than 18 fan blades. 
     
     
         11 . A gas turbine engine comprising:
 a fan section including a fan rotor and fan, the fan including a plurality of fan blades and with a low fan pressure ratio in operation of less than 1.45, the low fan pressure ratio measured across the fan blades alone;   a compressor section;   wherein the fan delivers a portion of air into a bypass duct, and a bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section, and the bypass ratio is greater than 10;   a gear train having a gear ratio greater than 2.3;   a combustor in fluid communication with the compressor section;   a turbine section in fluid communication with the combustor and the turbine section including a fan drive turbine driving the fan rotor through the gear train and a second turbine, wherein:
 the engine includes a power density greater than 1.5 lbf/in 3  and less than or equal to 5.5 lbf/in 3 , wherein power density is defined as Sea Level Takeoff Thrust in lbf produced divided by the volume of the turbine section in cubic inches; 
 the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio in operation greater than 5 in operation, wherein the fan drive turbine pressure ratio is a ratio of a pressure measured prior to the inlet as related to a pressure at the outlet prior to any exhaust nozzle; 
 the fan drive turbine has a first exit area and rotates at a first speed, the second turbine has a second exit area and rotates at a second speed, which is faster than the first speed, wherein a first performance quantity is defined as the product of the first speed squared and the first exit area, a second performance quantity is defined as the product of the second speed squared and the second exit area, and a performance ratio of the first performance quantity to the second performance quantity is between 0.5 and 1.5; and 
 the fan drive turbine has a number of fan drive turbine stages, and a ratio between the number of fan blades and the number of fan drive turbine stages is between 2.5 and 8.5. 
   
     
     
         12 . The gas turbine engine set forth in  claim 11 , wherein the second turbine is a two stage turbine. 
     
     
         13 . The gas turbine engine set forth in  claim 12 , wherein the performance ratio is greater than or equal to 0.8. 
     
     
         14 . The gas turbine engine set forth in  claim 13 , wherein the fan section has a low corrected fan tip speed less than 1150 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed divided by [(Tram° R)/(518.7° R)] 0.5 . 
     
     
         15 . The gas turbine engine set forth in  claim 14 , wherein the plurality of fan blades is less than 18 fan blades. 
     
     
         16 . The gas turbine engine set forth in  claim 14 , wherein the fan drive turbine includes at least one rotor having a live rim radius (r) and a bore width (W), and a ratio of r/W is between 4.65 and 5.55. 
     
     
         17 . The gas turbine engine set forth in  claim 16 , wherein the at least one rotor having a bore radius (R) and a live rim radius (r), and a ratio of r/R is between 2.00 and 2.30. 
     
     
         18 . The gas turbine engine set forth in  claim 11 , wherein the fan drive turbine includes at least one rotor having a live rim radius (r) and a bore width (W), and a ratio of r/W is between 4.65 and 5.55. 
     
     
         19 . The gas turbine engine set forth in  claim 11 , wherein the plurality of fan blades is less than 18 fan blades. 
     
     
         20 . The gas turbine engine set forth in  claim 11 , wherein the fan section has a low corrected fan tip speed less than 1150 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed divided by [(Tram° R)/(518.7° R)] 0.5 . 
     
     
         21 . The gas turbine engine set forth in  claim 20 , wherein the performance ratio is greater than or equal to 0.8. 
     
     
         22 . A gas turbine engine comprising:
 a fan section including a fan rotor and fan, the fan including a plurality of fan blades;   a gear train having a gear ratio greater than 2.3;   a compressor section;   a combustor in fluid communication with the compressor section;   a turbine section in fluid communication with the combustor and including a fan drive turbine and a second turbine, the fan drive turbine driving the fan rotor through the gear train wherein:
 the engine includes a power density greater than 1.5 lbf/in 3  and less than or equal to 5.5 lbf/in 3 , wherein power density is defined as Sea Level Takeoff Thrust in lbf produced divided by the volume of the turbine section in cubic inches; 
 the fan drive turbine has a first exit area and rotates at a first speed, the second turbine has a second exit area and rotates at a second speed, which is faster than the first speed, wherein a first performance quantity is defined as the product of the first speed squared and the first exit area, a second performance quantity is defined as the product of the second speed squared and the second exit area, and a performance ratio of the first performance quantity to the second performance quantity is between 0.5 and 1.5; and 
 the fan drive turbine has a number of fan drive turbine stages, and a ratio between the number of fan blades and the number of fan drive turbine stages is between 2.5 and 8.5. 
   
     
     
         23 . The gas turbine engine set forth in  claim 22 , wherein the fan section has a low corrected fan tip speed less than 1150 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed divided by [(Tram° R)/(518.7° R)] 0.5 . 
     
     
         24 . The gas turbine engine set forth in  claim 22 , wherein the fan section includes a low fan pressure ratio in operation of less than 1.45 at cruise at 0.8 Mach and 35,000 feet, the low fan pressure ratio measured across the fan blades alone, wherein the fan delivers a portion of air into a bypass duct, and a bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section, and the bypass ratio is greater than 10 at cruise at 0.8 Mach and 35,000. 
     
     
         25 . The gas turbine engine set forth in  claim 24 , wherein the plurality of fan blades is less than 18 fan blades. 
     
     
         26 . The gas turbine engine set forth in  claim 24 , wherein the performance ratio is greater than or equal to 0.8. 
     
     
         27 . The gas turbine engine set forth in  claim 26 , wherein the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio greater than 5 in operation, wherein the fan drive turbine pressure ratio is a ratio of a pressure measured prior to the inlet as related to a pressure at the outlet prior to any exhaust nozzle. 
     
     
         28 . The gas turbine engine set forth in  claim 22 , wherein the second turbine is a two stage turbine, the gear train is a planetary gear system, and the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio greater than 5 in operation, wherein the fan drive turbine pressure ratio is a ratio of a pressure measured prior to the inlet as related to a pressure at the outlet prior to any exhaust nozzle. 
     
     
         29 . The gas turbine engine set forth in  claim 28 , wherein the performance ratio is greater than or equal to 0.8. 
     
     
         30 . The gas turbine engine set forth in  claim 29 , wherein the fan drive turbine includes at least one rotor having a bore width (W) and a live rim radius (r), and a ratio of r/W is between 4.65 and 5.55.