Low noise compressor and turbine 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: 5500≤(number of blades×rotational speed)/60≤10000, the rotational speed being 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 gear reduction including an epicyclic gear train, wherein the gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine, 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 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 and the rotational speed being such that the following formula holds true for at least one of the blade rows of the low pressure compressor:   5500 Hz≤(number of blades×rotational speed)/60≤10000 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   the following formula holds true for at least one of the blade rows of the low pressure turbine:   (number of blades×rotational speed)/60<10000 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 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 less than or equal to 10000 for at least a plurality of the blade rows of the low pressure compressor. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number greater than or equal to 5500 for at least half of the blade rows of the low pressure compressor. 
     
     
         4 . The gas turbine engine as set forth in  claim 3 , wherein the formula results in a number greater than or equal to 5500 for at least one of the blade rows of the low pressure turbine. 
     
     
         5 . The gas turbine engine as set forth in  claim 4 , wherein the formula results in a number greater than or equal to 6000 for at least one of the blade rows of the low pressure compressor. 
     
     
         6 . 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 a plurality of the blade rows of the low pressure turbine. 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein the formula results in a number greater than or equal to 6000 for at least half of the blade rows of the low pressure compressor. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein the formula results in a number greater than or equal to 6000 for at least half of the blade rows of the low pressure turbine. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein the formula results in a number greater than or equal to 6000 for three of the blade rows of the low pressure turbine. 
     
     
         10 . The gas turbine engine as set forth in  claim 9 , wherein the formula results in a number less than or equal to 10000 for at least a majority of the blade rows of the low pressure turbine. 
     
     
         11 . 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 low pressure compressor. 
     
     
         12 . The gas turbine engine as set forth in  claim 11 , 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; and   wherein the fan has a low corrected fan tip speed of less than 1150 ft/second.   
     
     
         13 . The gas turbine engine as set forth in  claim 12 , wherein the gear reduction is a planetary gear system. 
     
     
         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 gear reduction including an epicyclic gear train, wherein the gear reduction is 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, 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 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 and the rotational speed being such that the following formula holds true for at least one of the blade rows of the low pressure compressor:   5500 Hz≤(number of blades×rotational speed)/60≤10000 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   the following formula holds true for at least one of the blade rows of the low pressure turbine:   (number of blades×rotational speed)/60≤10000 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 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 less than or equal to 10000 for at least a plurality of the blade rows of the low pressure compressor. 
     
     
         16 . The gas turbine engine as set forth in  claim 15 , wherein the formula results in a number greater than or equal to 5500 for at least one of the blade 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 for at least one of the blade rows of the low pressure compressor and for at least one of the blade rows of the low pressure turbine. 
     
     
         18 . The gas turbine engine as set forth in  claim 17 , wherein the formula results in a number greater than or equal to 6000 for at least half of the blade rows of the low pressure compressor. 
     
     
         19 . The gas turbine engine as set forth in  claim 18 , wherein the formula results in a number greater than or equal to 6000 for at least half of the blade rows of the low pressure turbine. 
     
     
         20 . The gas turbine engine as set forth in  claim 19 , wherein the formula results in a number less than or equal to 10000 for at least a plurality of the blade rows of the low pressure turbine. 
     
     
         21 . The gas turbine engine as set forth in  claim 20 , wherein the formula results in a number less than or equal to 7000 for only a majority of the blade rows of the low pressure compressor and for only a majority of the blade rows of the low pressure turbine. 
     
     
         22 . The gas turbine engine as set forth in  claim 20 , wherein the formula results in a number less than or equal to 10000 for at least a majority of the blade rows of the low pressure compressor and for at least a majority of the blade rows of the low pressure turbine. 
     
     
         23 . The gas turbine engine as set forth in  claim 22 , wherein the formula results in a number less than or equal to 7000 for at least a majority of the blade rows of the low pressure compressor and for at least a majority of the blade rows of the low pressure turbine, and the formula results in a number greater than or equal to 6000 for at least a majority of the blade rows of the low pressure compressor and for at least a majority of the blade rows of the low pressure turbine. 
     
     
         24 . The gas turbine engine as set forth in  claim 23 , 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.   
     
     
         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 gear reduction including an epicyclic gear train, wherein the gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine, 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 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 and the rotational speed being such that the following formula holds true for at least one of the blade rows of the low pressure compressor:   6000 Hz≤(number of blades×rotational speed)/60≤10000 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   the following formula holds true for at least one of the blade rows of the low pressure turbine:   (number of blades×rotational speed)/60≤10000 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 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 less than or equal to 10000 for at least a plurality of the blade rows of the low pressure compressor and for at least a plurality of the blade 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 5500 for at least half of the blade rows of the low pressure compressor and for at least half of the blade rows of the low pressure turbine. 
     
     
         28 . The gas turbine engine as set forth in  claim 27 , wherein the formula results in a number less than or equal to 7000 for at least a majority of the blade rows of the low pressure compressor and for at least a majority of the blade 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 for at least a majority of the blade rows of the low pressure compressor and for at least a majority of the blade 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, and the gear reduction is 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;   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.