Low noise compressor rotor for geared turbofan engine

Abstract

A gas turbine engine has a fan section including a fan, a compressor section including a low pressure compressor and a high pressure compressor, a turbine section including a low pressure turbine and a high pressure turbine, and a gear reduction. The low pressure compressor and the low pressure turbine have a number of blades in each of at least one of a plurality of blade rows. The blades are rotatable at least some of the time at a rotational speed in operation. The number of 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 turbine: (number of blades×rotational speed)/60s≥5500 Hz, and the rotational speed is in revolutions per minute.

What is claimed is: 
     
         1 . A gas turbine engine comprising:
 a fan section including a fan, and a low fan pressure ratio of less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a compressor section including a low pressure compressor and a high pressure compressor;   wherein the fan delivers air into a bypass duct defined within a nacelle, and a portion of air into the 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 the bypass ratio being greater than 10;   a turbine section including a low pressure turbine and a high pressure turbine;   wherein the low pressure turbine includes an inlet, an outlet and a pressure ratio of greater than 5, the pressure ratio being pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle;   a combustion section including a combustor between the high pressure compressor and the high pressure turbine;   a gear reduction including an epicyclic gear train, the gear reduction positioned between the fan on one side and the low pressure turbine on another side to drive the fan at a lower speed than the low pressure turbine in operation, and the epicyclic gear train having a gear reduction ratio of greater than 2.5:1;   wherein each of the low pressure compressor and the low pressure turbine includes a number of blades in each of at least one of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades in the at least one row and the rotational speed being such that the following formula holds true for the at least one row of the low pressure compressor:
   (the number of blades×the rotational speed)/60 sec≥5500 Hz;
 
   the 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 the gas turbine 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 holds true for at least a plurality of the rows of the low pressure compressor. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number less than or equal to 6000 Hz for at least one of the rows of the low pressure compressor. 
     
     
         4 . The gas turbine engine as set forth in  claim 3 , wherein the formula results in a number less than or equal to 6000 Hz for at least a majority of the rows of the low pressure compressor, and the formula results in a number greater than or equal to 5500 Hz for at least a majority of the rows of the low pressure compressor. 
     
     
         5 . The gas turbine engine as set forth in  claim 4 , further comprising:
 a core flowpath and a mid-turbine frame arranged between the low pressure turbine and the high pressure turbine, the mid-turbine frame having airfoils positioned in the core flowpath, and the mid-turbine frame supporting at least one bearing system;   wherein the low pressure compressor includes three stages, the low pressure turbine includes five stages, the high pressure turbine includes two stages, and the high pressure compressor includes eight stages; and   wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.   
     
     
         6 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number greater than or equal to 5500 Hz for at least a majority of the rows of the low pressure compressor. 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein the formula results in a number greater than or equal to 6000 Hz for at least one of the rows of the low pressure compressor. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein and the formula results in a number greater than or equal to 6000 Hz for at least a majority of the rows of the low pressure compressor. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein the formula results in a number greater than or equal to 5500 Hz for at least one of the rows of the low pressure turbine. 
     
     
         10 . The gas turbine engine as set forth in  claim 9 , wherein the formula results in a number greater than or equal to 5500 Hz for at least a majority of the rows of the low pressure turbine. 
     
     
         11 . The gas turbine engine as set forth in  claim 10 , wherein the formula results in a number greater than or equal to 6000 Hz for all of the rows of the low pressure compressor and for all of the rows of the low pressure turbine. 
     
     
         12 . The gas turbine engine as set forth in  claim 10 , wherein the formula results in a number greater than or equal to 6000 Hz for only a majority of the rows of the low pressure compressor and for only a majority of the rows of the low pressure turbine. 
     
     
         13 . The gas turbine engine as set forth in  claim 12 , further comprising:
 a core flowpath and a mid-turbine frame arranged between the low pressure turbine and the high pressure turbine, the mid-turbine frame having airfoils positioned in the core flowpath, and the mid-turbine frame supporting at least one bearing system;   wherein the gear reduction is a planetary gear system;   wherein the low pressure compressor is a three stage compressor, the low pressure turbine is a five stage turbine, the high pressure turbine is a two stage turbine, and the high pressure compressor is an eight stage compressor; and   wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.   
     
     
         14 . A gas turbine engine comprising:
 a fan section including a fan, and a low fan pressure ratio of less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a compressor section including a low pressure compressor and a high pressure compressor;   wherein the fan delivers air into a bypass duct defined within a nacelle, and a portion of air into the 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 the bypass ratio being greater than 10;   a turbine section including a low pressure turbine and a high pressure turbine;   wherein the low pressure turbine includes an inlet, an outlet and a pressure ratio of greater than 5, the pressure ratio being pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle;   a combustion section including a combustor between the high pressure compressor and the high pressure turbine;   a gear reduction including an epicyclic gear train, the gear reduction positioned intermediate the low pressure compressor and a shaft driven by the low pressure turbine such that a fan rotor of the fan section and the low pressure compressor are rotatable at a common speed in operation, and the epicyclic gear train having a gear reduction ratio of greater than 2.5:1;   wherein each of the low pressure compressor and the low pressure turbine includes a number of blades in each of at least one of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades in the at least one row and the rotational speed being such that the following formula holds true for the at least one row of the low pressure compressor:
   (the number of blades×the rotational speed)/60 sec≥5500 Hz;
 
   the 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 the gas turbine engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         15 . The gas turbine engine as set forth in  claim 14 , wherein the formula results in a number greater than or equal to 6000 Hz for at least a plurality of the rows of the low pressure compressor. 
     
     
         16 . The gas turbine engine as set forth in  claim 15 , wherein the formula does not hold true for any of the rows of the low pressure turbine. 
     
     
         17 . The gas turbine engine as set forth in  claim 16 , wherein the formula results in a number greater than or equal to 6000 Hz for only a majority of the rows of the low pressure compressor. 
     
     
         18 . The gas turbine engine as set forth in  claim 16 , wherein the formula results in a number greater than or equal to 6000 Hz for all of the rows of the low pressure compressor. 
     
     
         19 . The gas turbine engine as set forth in  claim 15 , wherein the formula results in a number greater than or equal to 5500 Hz for at least a plurality of the rows of the low pressure turbine. 
     
     
         20 . The gas turbine engine as set forth in  claim 19 , wherein the formula results in a number greater than or equal to 6000 Hz for at least one of the rows of the low pressure turbine. 
     
     
         21 . The gas turbine engine as set forth in  claim 20 , wherein the formula results in a number greater than or equal to 6000 Hz for at least a majority of the rows of the low pressure compressor and for at least a majority of the rows of the low pressure turbine. 
     
     
         22 . The gas turbine engine as set forth in  claim 21 , wherein the formula results in a number greater than or equal to 6000 Hz for all of the rows of the low pressure compressor, but for only a majority of the rows of the low pressure turbine. 
     
     
         23 . The gas turbine engine as set forth in  claim 21 , wherein the formula results in a number greater than or equal to 6000 Hz for all of the rows of the low pressure turbine, but for only a majority of the rows of the low pressure compressor. 
     
     
         24 . The gas turbine engine as set forth in  claim 21 , further comprising:
 a core flowpath and a mid-turbine frame arranged between the low pressure turbine and the high pressure turbine, the mid-turbine frame having airfoils positioned in the core flowpath, and the mid-turbine frame supporting at least one bearing system;   wherein the low pressure compressor includes three stages, the low pressure turbine includes a greater number of stages than the low pressure compressor, the high pressure turbine includes two stages, and the high pressure compressor includes a greater number of stages than the low pressure turbine;   wherein the fan has a low corrected fan tip speed of less than 1150 ft/second; and   wherein the formula results in a number greater than or equal to 5500 Hz for all of the rows of the low pressure compressor and for all of the rows of the low pressure turbine.   
     
     
         25 . A gas turbine engine comprising:
 a fan section including a fan, and a low fan pressure ratio of less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a compressor section including a low pressure compressor, an intermediate pressure compressor and a high pressure compressor;   wherein the fan delivers air into a bypass duct defined within a nacelle, and a portion of air into the 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 the bypass ratio being greater than 10;   a turbine section including a low pressure turbine, an intermediate pressure turbine and a high pressure turbine, the intermediate pressure turbine driving the intermediate pressure compressor, and the high pressure turbine driving the high pressure compressor;   wherein the low pressure turbine includes an inlet, an outlet and a pressure ratio of greater than 5, the pressure ratio being pressure measured prior to the inlet as related to pressure at the outlet prior to an exhaust nozzle;   a combustion section including a combustor between the high pressure compressor and the high pressure turbine;   a gear reduction including an epicyclic gear train, the gear reduction positioned between the fan on one side and the low pressure turbine on another side to drive the fan at a lower speed than the low pressure turbine in operation, and the epicyclic gear train having a gear reduction ratio of greater than 2.5:1;   wherein each of the low pressure compressor and the low pressure turbine includes a number of blades in each of at least one of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades in the at least one row and the rotational speed being such that the following formula holds true for the at least one row of the low pressure compressor:
   (the number of blades×the rotational speed)/60 sec≥5500 Hz;
 
   the 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 the gas turbine engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         26 . The gas turbine engine as set forth in  claim 25 , wherein the formula results in a number greater than or equal to 5500 Hz for at least one of the rows of the low pressure turbine. 
     
     
         27 . The gas turbine engine as set forth in  claim 26 , wherein the formula results in a number greater than or equal to 6000 Hz for at least a plurality of the rows of the low pressure compressor. 
     
     
         28 . The gas turbine engine as set forth in  claim 27 , wherein the formula results in a number greater than or equal to 6000 Hz for at least a plurality of the rows of the low pressure turbine. 
     
     
         29 . The gas turbine engine as set forth in  claim 28 , wherein the formula results in a number greater than or equal to 6000 Hz for at least a majority of the rows of the low pressure compressor and for at least a majority of the rows of the low pressure turbine. 
     
     
         30 . The gas turbine engine as set forth in  claim 29 , further comprising:
 a core flowpath and a mid-turbine frame arranged between the low pressure turbine and the high pressure turbine, the mid-turbine frame having airfoils positioned in the core flowpath, and the mid-turbine frame supporting at least one bearing system;   wherein the gear reduction is a planetary gear system;   wherein the low pressure compressor includes three stages, the low pressure turbine includes five stages, the high pressure turbine includes two stages, and the high pressure compressor includes eight stages; and   wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.