Three spool turbofan engine with low noise intermediate turbine rotor

Abstract

A turbine section including a high pressure turbine, an intermediate pressure turbine and a fan drive turbine, the fan drive turbine driving a gear reduction to in turn drive a fan, and effecting a reduction in the speed of the fan relative to an input speed from the fan drive turbine and said high pressure turbine driving a high pressure compressor, and the intermediate pressure turbine driving a low pressure compressor, with the intermediate pressure turbine having a number of turbine blades in at least one row, and the turbine blades operating at least some of the time at a rotational speed, and the number of turbine 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 intermediate pressure turbine: (number of blades×speed)/60≥5500 Hz.

1 . A gas turbine engine comprising:
 a turbine section including a high pressure turbine, an intermediate pressure turbine and a fan drive turbine;   said fan drive turbine driving a gear reduction to in turn drive a fan, and effecting a reduction in the speed of the fan relative to an input speed from said fan drive turbine;   said high pressure turbine driving a high pressure compressor;   said intermediate pressure turbine driving a low pressure compressor;   said intermediate pressure turbine having a number of turbine blades greater than one in at least one row, and said turbine blades operating at least some of the time at an intermediate pressure turbine rotational speed, and said number of turbine 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 intermediate pressure turbine
   (number of blades×intermediate pressure turbine rotational speed in revolutions per minute)/60≥5500 Hz; and
 
   wherein said gas turbine engine is rated to produce 15,000 lbs. of thrust or more, and said rotational speed being an approach speed.   
     
     
         2 . The gas turbine engine as set forth in  claim 1 , wherein the formula results in a number greater than or equal to 6000 Hz. 
     
     
         3 . The gas turbine engine as set forth in  claim 2 , wherein the formula holds true for the majority of blade rows of the intermediate pressure turbine. 
     
     
         4 . The gas turbine engine as set forth in  claim 1 , wherein the formula holds true for the majority of blade rows of the intermediate pressure turbine. 
     
     
         5 . The gas turbine engine as set forth in  claim 1 , wherein the fan delivers air into a bypass duct and into a compressor section as core air, and a bypass ratio of the volume of air delivered into the bypass duct compared to the volume of air delivered as core air being greater than or equal to 6. 
     
     
         6 . The gas turbine engine as set forth in  claim 5 , wherein bypass ratio is greater than or equal to  10 . 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein a gear ratio is greater than or equal to 2.3. 
     
     
         8 . The gas turbine engine as set forth in  claim 1 , wherein a gear ratio is greater than or equal to 2.3. 
     
     
         9 . A method of designing a gas turbine engine comprising the steps of:
 including a gear reduction between a fan drive turbine and a fan, a high pressure turbine driving a high pressure compressor downstream of said fan, and including an intermediate pressure turbine intermediate said high pressure turbine and said fan drive turbine, said intermediate pressure turbine driving a low pressure compressor which is intermediate said fan and said high pressure compressor; and   selecting a number of blades in at least one row of the intermediate pressure turbine, in combination with a rotational speed in revolutions per minute of the intermediate pressure turbine, such that the following formula holds true for said at least one row of the intermediate pressure turbine
   (number of blades×speed)/60≥5500 Hz.
 
   
     
     
         10 . The method of designing a gas turbine engine as set forth in  claim 9 , wherein the formula results in a number greater than or equal to 6000. 
     
     
         11 . The method of designing a gas turbine engine as set forth in  claim 10 , wherein said gas turbine engine is rated to produce 15,000 pounds of thrust or more. 
     
     
         12 . The method as set forth in  claim 10 , wherein the formula holds true for the majority of the blade rows of the fan drive turbine. 
     
     
         13 . The method as set forth in  claim 9 , wherein said rotational speed being an approach speed. 
     
     
         14 . The method as set forth in  claim 9 , wherein said rotational speed being an approach speed. 
     
     
         15 . The method as set forth in  claim 9 , wherein said gas turbine engine is rated to produce 15,000 lbs. of thrust or more. 
     
     
         16 . A turbine module comprising:
 a fan drive rotor, a high pressure turbine, and an intermediate pressure turbine; and   said intermediate pressure turbine having a first blade row that includes a number of blades, the first blade row being capable of rotating at a rotational speed, so that when measuring said rotational speed in revolutions per minute:
   (number of blades×said rotational speed)/60≥5500 Hz.
 
   
     
     
         17 . The turbine module as set forth in  claim 16 , wherein the formula results in a number greater than or equal to 6000. 
     
     
         18 . The turbine module as set forth in  claim 17 , wherein the formula holds true for the majority of blade rows of the fan drive rotor. 
     
     
         19 . The turbine module as set forth in  claim 16 , wherein said rotational speed being an approach speed. 
     
     
         20 . The turbine module as set forth in  claim 16 , wherein the formula holds true for the majority of blade rows of the fan drive rotor.