Low noise compressor rotor for geared turbofan engine

Abstract

A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor, and having a gear reduction ratio of greater than 2.5:1. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor: (number of blades×rotational speed)/60 sec≧5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine also disclosed.

1 . A gas turbine engine comprising:
 a fan, a turbine section having a fan drive turbine, and a compressor section having a first compressor;   a gear reduction positioned between said fan on one side and said fan drive turbine on another side, the gear reduction including an epicycle gear train having a gear reduction ratio of greater than 2.5:1;   said first compressor having a number of compressor blades in at least one of a plurality of rows of said first compressor, and said blades rotatable at least some of the time at a rotational speed in operation, and said number of compressor blades in said at least one row and said rotational speed being such that the following formula holds true for said at least one row of the first compressor:
   (said number of blades×said rotational speed)/60 sec≧5500 Hz;
 
   said rotational speed being an approach speed in revolutions per minute, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations;   wherein a pressure ratio across the fan drive turbine being greater than 5; and   wherein the gas turbine engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         2 - 3 . (canceled) 
     
     
         4 . The gas turbine engine as set forth in  claim 1 , wherein said fan delivers air into a bypass duct, and a portion of air into said compressor section, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section, and said bypass ratio being greater than 10. 
     
     
         5 - 6 . (canceled) 
     
     
         7 . The gas turbine engine as set forth in  claim 4 , wherein said fan comprises at least one fan blade, with a low fan pressure ratio of less than 1.45, the low fan pressure ratio measured across the fan blade alone. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein said turbine section includes a higher pressure turbine and a lower pressure turbine, and said fan drive turbine being said lower pressure turbine. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein said first compressor is a lower pressure compressor, and said higher pressure turbine drives a higher pressure compressor. 
     
     
         10 . The gas turbine engine as set forth in  claim 9 , wherein the gear reduction is positioned intermediate the fan drive turbine and said first compressor. 
     
     
         11 . The gas turbine engine as set forth in  claim 7 , wherein the formula does not hold true for all of the rows of said first compressor. 
     
     
         12 . The gas turbine engine as set forth in claim  77 , wherein the formula holds true for at least a majority of the rows of said first compressor. 
     
     
         13 . (canceled) 
     
     
         14 . The gas turbine engine as set forth in  claim 34 , wherein the fan has a low corrected fan tip speed of less than 1150 ft/second. 
     
     
         15 . The gas turbine engine as set forth in  claim 14 , further comprising a core flowpath and a mid-turbine frame arranged between the fan drive turbine and a second turbine and having airfoils positioned in the core flowpath, the mid-turbine frame supporting at least one bearing system, and wherein the formula results in a number greater than or equal to 6000 Hz for the at least one compressor row. 
     
     
         16 . A method of designing a gas turbine engine comprising the steps of:
 including a turbine section having a fan turbine for driving a fan through a gear reduction, the gear reduction positioned between the fan on one side and the fan turbine on another side and having a gear reduction ratio of greater than 2.5:1;   providing a first compressor including a plurality of blade rows;   selecting a combination of a number of blades in at least one row of the first compressor and a rotational speed of the first compressor in operation, for producing noise frequencies that are of less concern to humans;   said selecting step including determining the number of blades in the at least one row of the first compressor and determining a rotational speed of the first compressor in operation such that the following formula holds true for said at least one row of the first compressor:
   (said number of blades×said rotational speed)/60 sec≧5500 Hz;
 
   said rotational speed being 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 a pressure ratio across the fan turbine being greater than 5, and the gas turbine engine is rated to produce 15,000 pounds of thrust or more. 
   
     
     
         17 - 18 . (canceled) 
     
     
         19 . The method as set forth in  claim 16 , wherein said first compressor is a lower pressure compressor, and said higher pressure turbine driving a higher pressure compressor. 
     
     
         20 . The method as set forth in  claim 19 , wherein:
 the fan comprises at least one fan blade, with a low fan pressure ratio of less than 1.45, the low fan pressure ratio measured across the fan blade alone; and   the fan delivers air into a bypass duct, and a portion of air into said first compressor, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into said first compressor, and said bypass ratio being greater than 10.   
     
     
         21 . (canceled) 
     
     
         22 . The method as set forth in  claim 20 , wherein the formula does not hold true for all of the rows of the first compressor. 
     
     
         23 . The method as set forth in  claim 35 , wherein the formula holds true for at least a majority of the rows of the first compressor. 
     
     
         24 . The method as set forth in  claim 23 , wherein the formula results in a number greater than or equal to 6000 Hz for at least one compressor row. 
     
     
         25 . A gas turbine engine comprising:
 a fan, a turbine section having a fan drive turbine, and a compressor section having a first compressor;   a gear reduction positioned between said fan on one side and said fan drive turbine on another side and having a gear reduction ratio of greater than 2.5:1;   said first compressor having a number of compressor blades in at least one of a plurality of rows of said first compressor, and said blades rotatable at least some of the time at a rotational speed in operation, and said number of compressor blades in said at least one row and said rotational speed being such that the following formula holds true for said at least one row of the first compressor:
   5500 Hz≦(said number of blades×said rotational speed)/60 sec≦6000 Hz;
 
   said rotational speed being 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 a pressure ratio across the fan drive turbine being greater than 5.   
     
     
         26 . (canceled) 
     
     
         27 . The gas turbine engine as set forth in  claim 25 , wherein the formula does not hold true for all of the rows of the first compressor. 
     
     
         28 . (canceled) 
     
     
         29 . The gas turbine engine as set forth in  claim 25 , wherein:
 said fan delivers air into a bypass duct, and a portion of air into said compressor section, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section, and said bypass ratio being greater than 10.   
     
     
         30 . The gas turbine engine as set forth in  claim 38 , wherein the formula holds true for at least a majority of the rows of the first compressor. 
     
     
         31 . The gas turbine engine as set forth in  claim 11 , wherein the formula holds true for at least a majority of the rows of said first compressor. 
     
     
         32 . The gas turbine engine as set forth in  claim 12 , wherein the formula holds true for all of the rows of said first compressor. 
     
     
         33 . The gas turbine engine as set forth in  claim 12 , wherein the formula results in a number greater than or equal to 6000 Hz for at least one of the majority of compressor rows. 
     
     
         34 . The gas turbine engine as set forth in  claim 33 , wherein the formula results in a number greater than or equal to 6000 Hz for each of the majority of compressor rows. 
     
     
         35 . The method as set forth in  claim 22 , wherein the formula holds true for at least a plurality of the rows of the first compressor. 
     
     
         36 . The method as set forth in  claim 24 , wherein the formula results in a number greater than or equal to 6000 Hz for more than one row of the first compressor. 
     
     
         37 . The gas turbine engine as set forth in  claim 29 , wherein the formula holds true for more than one row of the first compressor. 
     
     
         38 . The gas turbine engine as set forth in  claim 37 , wherein the fan comprises at least one fan blade, with a low fan pressure ratio of less than 1.45, the low fan pressure ratio measured across the fan blade alone. 
     
     
         39 . The gas turbine engine as set forth in  claim 30 , wherein the formula holds true for all of the rows of said first compressor.