Low noise turbine for geared turbofan engine

Abstract

A method of designing a gas turbine engine comprises the steps of including a fan section with a fan. A turbine section is included having a first turbine and a high pressure turbine. A gear reduction is included between the fan and the first turbine, the gear reduction being configured to receive an input from the first turbine and to turn the fan at a lower speed than the first turbine in operation. The first turbine is designed to include a number of turbine blades in each of a plurality of rows of the first turbine, the first turbine blades operating at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least one of the blade rows of the first turbine: (number of blades×speed)/60≥5500.

What is claimed is: 
     
         1 . A method of designing a gas turbine engine comprising the steps of:
 including a fan section with a fan, the fan including at least one fan blade;   the fan section designed to achieve a low fan pressure ratio less than about 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   further including a turbine section having a first turbine and a second turbine;   further including a gear reduction between the fan and the first turbine, the gear reduction including an epicycle gear train having a gear reduction ratio of greater than about 2.5:1, and the gear reduction being configured to receive an input from the first turbine and to turn the fan at a lower speed than the first turbine in operation;   the first turbine designed to achieve a pressure ratio greater than about 5:1, the first turbine including an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the first turbine being a ratio of the inlet pressure to the outlet pressure; and   the first turbine designed to further include a number of turbine blades in each of a plurality of rows of the first turbine, the first turbine blades operating at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least one of the blade rows of the first turbine: (number of blades×speed)/60≥5500;   wherein the rotational speed is an approach speed in revolutions per minute, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations; and   wherein the gas turbine engine is designed to produce 15,000 pounds of thrust or more.   
     
     
         2 . The method as recited in  claim 1 , wherein the formula results in a number greater than 6000. 
     
     
         3 . The method as recited in  claim 2 , wherein the formula results in a number less than or equal to about 10000. 
     
     
         4 . The method as recited in  claim 3 , wherein the formula results in a number less than 7000. 
     
     
         5 . The method as recited in  claim 1 , wherein the formula holds true for a majority of the blade rows of the first turbine. 
     
     
         6 . The method as recited in  claim 5 , wherein the formula results in a number greater than 6000. 
     
     
         7 . The method as recited in  claim 6 , wherein the formula results in a number less than or equal to about 10000. 
     
     
         8 . The method as recited in  claim 7 , wherein the formula results in a number less than 7000. 
     
     
         9 . The method as recited in  claim 5 , further including a mid-turbine frame arranged between the second turbine and the first turbine. 
     
     
         10 . The method as recited in  claim 9 , further including a compressor section configured to drive air along core flowpath, and a plurality of bearing systems configured to support the first turbine and the second turbine, wherein the mid-turbine frame includes airfoils positioned in the core flowpath and is configured to support at least one of the bearing systems. 
     
     
         11 . The method as recited in  claim 10 , wherein the second turbine has two stages. 
     
     
         12 . The method as recited in  claim 5 , further including a first compressor, and a shaft configured to be driven by the first turbine, the gear reduction arranged intermediate the first compressor and the shaft. 
     
     
         13 . The method as recited in  claim 12 , wherein the second turbine has two stages. 
     
     
         14 . The method as recited in  claim 5 , further designing the engine to achieve a bypass ratio greater than ten (10), and wherein the fan is designed to have a low corrected fan tip speed less than about 1150 ft/second, wherein the low corrected fan tip speed is an actual fan tip speed in ft/second at an ambient temperature divided by [(Tambient ° R)/(518.7° R)] 0.5 . 
     
     
         15 . The method as recited in  claim 14 , wherein the fan section is designed for cruise. 
     
     
         16 . The method as recited in  claim 1 , wherein the formula holds true for all of the blade rows of the first turbine. 
     
     
         17 . The method as recited in  claim 16 , wherein the formula results in a number greater than 6000. 
     
     
         18 . The method as recited in  claim 17 , wherein the formula results in a number less than or equal to about 10000. 
     
     
         19 . The method as recited in  claim 18 , wherein the formula results in a number less than 7000. 
     
     
         20 . The method as recited in  claim 1 , wherein the formula does not hold true for all of the blade rows of the first turbine.