Low noise compressor and turbine for geared turbofan engine

Abstract

A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine rotor, a compressor rotor, and a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor. Each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate 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 a plurality of the blade rows of the first turbine rotor: (number of blades×rotational speed)/60≧5500, and the rotational speed being an approach speed in revolutions per minute, and the following formula holds true for at least a plurality of the blade rows of the compressor rotor: (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a fan;   a low fan pressure ratio less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a bypass ratio of greater than 10;   a turbine section having a first turbine including a first turbine rotor, wherein the first turbine includes a pressure ratio greater than 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;   a compressor rotor;   a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor, the gear reduction having a gear reduction ratio of greater than 2.5:1; and   wherein each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate 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 a plurality of the blade rows of the first turbine rotor:   (number of blades×rotational speed)/60≧5500, the rotational speed being an approach speed in revolutions per minute; and   the following formula holds true for at least a plurality of the blade rows of the compressor rotor:   (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute; and   wherein the engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         2 . The gas turbine engine as set forth in  claim 1 , wherein the formula results in a number greater than 6000 for at least a majority of the blade rows of the first turbine rotor. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number less than or equal to 10000 for at least one of the plurality of blade rows of the first turbine rotor. 
     
     
         4 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number less than or equal to 10000 for all blade rows of the compressor rotor. 
     
     
         5 . The gas turbine engine as set forth in  claim 4 , wherein the formula results in a number less than 7000 for at least one blade row of the compressor rotor. 
     
     
         6 . The gas turbine engine as set forth in  claim 5 , wherein the formula results in a number greater than 6000 for at least one blade row of the compressor rotor. 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein the formula results in a number less than 7000 for at least one blade row of the first turbine rotor. 
     
     
         8 . The gas turbine engine as set forth in  claim 5 , wherein the formula results in a number less than or equal to 10000 for at least one blade row of the first turbine rotor. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein the formula results in a number less than or equal to 10000 for at least a plurality of blade rows of the first turbine rotor. 
     
     
         10 . A gas turbine engine comprising:
 a fan;   a low fan pressure ratio less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a bypass ratio of greater than 10;   a turbine section having a first turbine including a first turbine rotor, wherein the first turbine includes a pressure ratio greater than 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;   a compressor rotor;   a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor, the gear reduction having a gear reduction ratio of greater than 2.5:1; and   wherein each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate 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 a majority of the blade rows of the first turbine rotor:   (number of blades×rotational speed)/60≧6000, the rotational speed being an approach speed in revolutions per minute;   the following formula holds true for at least a plurality of the blade rows of the compressor rotor:   (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute; and   wherein the engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         11 . The gas turbine engine as set forth in  claim 10 , wherein the formula results in a number greater than or equal to 6000 for three blade rows of the first turbine rotor. 
     
     
         12 . The gas turbine engine as set forth in  claim 11 , wherein the formula results in a number less than or equal to 10000 for at least a majority of the blade rows of the first turbine rotor. 
     
     
         13 . The gas turbine engine as set forth in  claim 12 , wherein the formula results in a number less than or equal to 10000 for all blade rows of the first turbine rotor. 
     
     
         14 . The gas turbine engine as set forth in  claim 13 , wherein the formula results in a number greater than or equal to 6000 for all blade rows of the first turbine rotor, and the formula results in a number less than or equal to 7000 for at least one blade row of the compressor rotor. 
     
     
         15 . The gas turbine engine as set forth in  claim 14 , wherein the formula results in a number greater than or equal to 5500 for at least one blade row of the compressor rotor. 
     
     
         16 . The gas turbine engine as set forth in  claim 12 , wherein the formula results in a number greater than or equal to 6000 for at least a majority of the blade rows of the first turbine rotor, and the formula results in a number less than or equal to 7000 for at least a plurality of blade rows of the compressor rotor. 
     
     
         17 . The gas turbine engine as set forth in  claim 10 , wherein the formula results in a number less than or equal to 10000 for at least a majority of the blade rows of the first turbine rotor. 
     
     
         18 . The gas turbine engine as set forth in  claim 17 , wherein the formula results in a number greater than 6000 for all blade rows of the first turbine rotor. 
     
     
         19 . The gas turbine engine as set forth in  claim 18 , wherein the formula results in a number less than 7000 for at least one blade row of the first turbine rotor. 
     
     
         20 . The gas turbine engine as set forth in  claim 19 , wherein the formula results in a number less than 7000 for at least a plurality of blade rows of the first turbine rotor. 
     
     
         21 . The gas turbine engine as set forth in  claim 19 , wherein the formula results in a number less than 7000 for at least one blade row of the compressor rotor. 
     
     
         22 . The gas turbine engine as set forth in  claim 21 , wherein the formula results in a number less than 7000 for at least a plurality of blade rows of the compressor rotor. 
     
     
         23 . The gas turbine engine as set forth in  claim 22 , wherein the formula results in a number less than or equal to 10000 for all blade rows of the first turbine rotor. 
     
     
         24 . The gas turbine engine as set forth in  claim 18 , wherein the formula results in a number less than or equal to 10000 for at least a majority of the blade rows of the compressor rotor. 
     
     
         25 . The gas turbine engine as set forth in  claim 24 , wherein the formula results in a number less than or equal to 10000 for all blade rows of the compressor rotor. 
     
     
         26 . The gas turbine engine as set forth in  claim 25 , wherein the formula results in a number less than 7000 for at least one blade row of the compressor rotor. 
     
     
         27 . The gas turbine engine as set forth in  claim 26 , wherein the formula results in a number less than 7000 for at least one blade row of the first turbine rotor. 
     
     
         28 . A gas turbine engine comprising:
 a fan;   a low fan pressure ratio less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a bypass ratio of greater than 10;   a turbine section having a first turbine including a first turbine rotor, wherein the first turbine includes a pressure ratio greater than 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;   a compressor rotor;   a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor, the gear reduction having a gear reduction ratio of greater than 2.5:1; and   wherein each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate 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 all blade rows of the first turbine rotor:   (number of blades×rotational speed)/60≧6000, the rotational speed being an approach speed in revolutions per minute;   the following formula holds true for at least a plurality of the blade rows of the first turbine rotor:   (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute;   the following formula holds true for at least a majority of the blade rows of the compressor rotor:   (number of blades×rotational speed)/60≦10000, the rotational speed being an approach speed in revolutions per minute; and   wherein the engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         29 . The gas turbine engine as set forth in  claim 28 , wherein the formula results in a number less than 10000 for at least a plurality of blade rows of the first turbine rotor. 
     
     
         30 . The gas turbine engine as set forth in  claim 29 , wherein the formula results in a number less than 7000 for at least a majority of the blade rows of the compressor rotor.