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, a compressor, and a gear reduction positioned between the fan and the first turbine. Each of the compressor and the first 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 the plurality of the blade rows of the first turbine, but does not hold true for any of the blade rows of the compressor rotor: (number of blades×rotational speed)/60≧5500, and the rotational speed being an approach speed in revolutions per minute.

1 . A gas turbine engine comprising:
 a fan section including a fan, the fan including at least one fan blade;   a low fan pressure ratio less than about 1.45, wherein the low fan pressure ratio is measured across a fan blade alone;   a compressor rotationally coupled to the fan;   a turbine section having a first turbine and a second turbine;   a gear reduction positioned between the first turbine on one side and the compressor and fan on another side, the gear reduction including an epicycle gear train having a gear reduction ratio of greater than 2.5:1, and the first turbine positioned in a driving relationship with the gear reduction;   the first turbine including 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;   each of the compressor and the first 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 more than one of the blade rows of the first turbine, but does not hold true for any of the blade rows of the compressor:
   (number of blades×rotational speed)/60≧5500,
 
   
       and
 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 results in a number less than or equal to about 10000 for at least one of the blade rows of the first turbine, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein the following formula holds true for at least one of the blade rows of the first turbine:
   (number of blades×rotational speed)/60≦10000,
   the rotational speed being a cruise speed in revolutions per minute.   
     
     
         4 . The gas turbine engine as set forth in  claim 2 , wherein the formula results in a number less than or equal to 7000 Hz for at least one of the blade rows of the first turbine, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         5 . The gas turbine engine as set forth in  claim 1 , further comprising a bypass ratio greater than ten (10). 
     
     
         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 of the blade rows of the first turbine, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein the fan has a low corrected fan tip speed less than 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 . 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , further comprising a core flowpath and a mid-turbine frame having airfoils positioned in the core flowpath, the mid-turbine frame supporting at least one bearing system. 
     
     
         9 . The gas turbine engine as set forth in  claim 8 , wherein the second turbine has two stages, and the epicyclic gear train is a planetary gear system. 
     
     
         10 . The gas turbine engine as set forth in  claim 9 , wherein the following formula holds true for at least one of the blade rows of the first turbine:
   (number of blades×rotational speed)/60≧5500,
   the rotational speed being a cruise speed in revolutions per minute.   
     
     
         11 . A gas turbine engine comprising:
 a fan;   a turbine section including a first turbine;   a compressor;   a gear reduction positioned between the fan relative to an input speed from the first turbine; and   wherein each of the compressor and the first 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 a majority of the blade rows of the first turbine, and at least one of the blade rows of the compressor:
   (number of blades×rotational speed)/60≧5500, and
 
   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 engine is rated to produce 15,000 pounds of thrust or more.   
     
     
         12 . The gas turbine engine as set forth in  claim 11 , wherein the formula holds true for less than half of the blade rows of the compressor. 
     
     
         13 . The gas turbine engine as set forth in  claim 12 , wherein the formula holds true for all of the blade rows of the first turbine. 
     
     
         14 . The gas turbine engine as set forth in  claim 11 , wherein the following formula holds true for at least one of the blade rows of the first turbine:
   (number of blades×rotational speed)/60≧5500,
   the rotational speed being a cruise speed in revolutions per minute.   
     
     
         15 . The gas turbine engine as set forth in  claim 11 , further comprising a bypass ratio greater than ten (10), the gear reduction includes an epicycle gear train having a gear reduction ratio of greater than 2.5:1, and the first turbine is in a driving relationship with the compressor. 
     
     
         16 . The gas turbine engine as set forth in  claim 15 , further comprising a low fan pressure ratio less than 1.45, wherein the low fan pressure ratio is measured across a fan blade alone. 
     
     
         17 . The gas turbine engine as set forth in  claim 16 , 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. 
     
     
         18 . The gas turbine engine as set forth in  claim 17 , wherein the fan has a low corrected fan tip speed less than 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 . 
     
     
         19 . The gas turbine engine as set forth in  claim 17 , wherein the formula results in a number greater than or equal to 5500 but less than or equal to 10000 for at least one of the blade rows of the compressor, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         20 . The gas turbine engine as set forth in  claim 19 , wherein the compressor and the fan rotate at a common speed in operation. 
     
     
         21 . The gas turbine engine as set forth in  claim 20 , further comprising a core flowpath and a mid-turbine frame having airfoils positioned in the core flowpath. 
     
     
         22 . The gas turbine engine as set forth in  claim 11 , wherein the formula holds true for all of the blade rows of the first turbine, and results in a number greater than or equal to 5500 but less than or equal to 7000 for at least one of the blade rows of the compressor, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         23 . A gas turbine engine comprising:
 a fan;   a turbine section including a first turbine;   a compressor;   a gear reduction positioned between the fan relative to an input speed from the first turbine; and   wherein each of the compressor and the first 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 all of the blade rows of the first turbine, and all of the blade rows of the compressor:
   (number of blades×rotational speed)/60≧5500,
 
   
       and
 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. 
 
     
     
         24 . The gas turbine engine as set forth in  claim 23 , wherein the formula results in a number less than or equal to about 10000 for at least one of the blade rows of the first turbine, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         25 . The gas turbine engine as set forth in  claim 23 , 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. 
     
     
         26 . The gas turbine engine as set forth in  claim 25 , further comprising a low fan pressure ratio less than about 1.45, wherein the low fan pressure ratio is measured across a fan blade alone. 
     
     
         27 . The gas turbine engine as set forth in  claim 26 , wherein the gear reduction has a gear reduction ratio of greater than about 2.5:1. 
     
     
         28 . The gas turbine engine as set forth in  claim 27 , wherein the formula results in a number greater than or equal to 5500 but less than or equal to about 7000 for at least one of the blade rows of the compressor, taken at an approach certification point as defined in Part 36 of the Federal Airworthiness Regulations. 
     
     
         29 . The gas turbine engine as set forth in  claim 23 , wherein the gas turbine engine is rated to produce 15,000 pounds of thrust or more, and wherein the fan has 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 . 
     
     
         30 . The gas turbine engine as set forth in  claim 23 , wherein the following formula holds true for at least one of the blade rows of the first turbine:
   (number of blades×rotational speed)/60≧5500,
   the rotational speed being a cruise speed in revolutions per minute.