Geared turbofan architecture for regional jet aircraft

Abstract

A gas turbine engine comprises an upstream compressor rotor and a downstream compressor rotor, each having a first stage blade row and a tip speed defined at the first stage blade row. A ratio of the tip speed of the first stage of the downstream compressor rotor to the tip speed of the first stage in the upstream compressor rotor is equal to or above about 1.18.

1 . A gas turbine engine comprising:
 an upstream compressor rotor and a downstream compressor rotor, each of said compressor rotors having a first stage blade row and a tip speed defined at said first stage blade row, with a ratio of said tip speed of said downstream compressor rotor first stage blade row to the tip speed of said upstream compressor rotor first stage blade row being equal to or above about 1.18; and   said gas turbine engine is designed for use on a regional aircraft defined as having at least one of a single aisle between passenger sections or a total flight length of less than 3,000 nautical miles.   
     
     
         2 . The gas turbine engine section as set forth in  claim 1 , wherein said upstream compressor rotor having an upstream number of stages and said downstream compressor rotor having a downstream number of stages with a ratio of said downstream number of stages to said upstream number of stages being greater than or equal to about 1.6 and less than or equal to about 3.4. 
     
     
         3 . The gas turbine engine section as set forth in  claim 2 , wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38. 
     
     
         4 . The gas turbine engine as set forth in  claim 3 , wherein a fan rotor having less than or equal to 26 blades. 
     
     
         5 . The gas turbine engine as set forth in  claim 3 , wherein a turbine section including a lower pressure turbine driving said upstream compressor rotor, said lower pressure turbine having three or four stages. 
     
     
         6 . The gas turbine engine as set forth in  claim 5 , wherein said lower pressure turbine also driving said fan rotor. 
     
     
         7 . The gas turbine engine as set forth in  claim 6 , wherein said lower pressure turbine driving said fan rotor through a gear reduction. 
     
     
         8 . The gas turbine engine as set forth in  claim 7 , wherein a gear ratio of said gear reduction being greater than or equal to about 2.6. 
     
     
         9 . The gas turbine engine as set forth in  claim 7 , wherein a pressure ratio across said fan rotor is less than or equal to about 1.45. 
     
     
         10 . The gas turbine engine as set forth in  claim 7 , wherein said gear reduction is a planet-type epicyclic gearbox. 
     
     
         11 . The gas turbine engine as set forth in  claim 10 , wherein said gas turbine engine is designed for use on a single aisle aircraft. 
     
     
         12 . The gas turbine engine as set forth in  claim 1 , wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38. 
     
     
         13 . The gas turbine engine as set forth in  claim 12 , wherein said gas turbine engine is designed for use on a single aisle aircraft. 
     
     
         14 . The gas turbine engine as set forth in  claim 12 , wherein said compression ratio being less than or equal to about 55. 
     
     
         15 . The gas turbine engine as set forth in  claim 1 , wherein a turbine section including a lower pressure turbine driving said upstream compressor rotor, said lower pressure turbine having three or four stages. 
     
     
         16 . The gas turbine engine as set forth in  claim 15 , wherein said lower pressure turbine also driving a fan rotor. 
     
     
         17 . The gas turbine engine as set forth in  claim 16 , wherein said fan rotor is driven through a gear reduction. 
     
     
         18 . The gas turbine engine as set forth in  claim 17 , wherein a gear ratio of said gear reduction being greater than or equal to about 2.6. 
     
     
         19 . The gas turbine engine as set forth in  claim 18 , wherein the gear ratio of said gear reduction being less than or equal to about 3.3. 
     
     
         20 . The gas turbine engine as set forth in  claim 17 , wherein a pressure ratio across said fan rotor is less than or equal to about 1.45. 
     
     
         21 . The gas turbine engine as set forth in  claim 1 , wherein said gas turbine engine is designed for use on a single aisle aircraft. 
     
     
         22 . The gas turbine engine as set forth in  claim 1 , wherein said ratio of said tip speed of the first stage of the downstream compressor rotor to the tip speed of the first stage in the upstream compressor rotor being and equal to or below about 1.43. 
     
     
         23 . The gas turbine engine as set forth in  claim 1 , wherein a fan rotor has less than or equal to 26 blades. 
     
     
         24 . A gas turbine engine comprising:
 an upstream compressor rotor and a downstream compressor rotor, each of said compressor rotors having a first stage blade row and a tip speed defined at said first stage blade row, with a ratio of said tip speed of said downstream compressor rotor first stage blade row to the tip speed of said upstream compressor rotor first stage blade row being equal to or below about 1.43, and said ratio being greater than or equal to about 1.18, and a fan rotor is included upstream of said upstream compressor rotor, and said fan rotor being driven by a turbine rotor through a gear reduction.   
     
     
         25 . The gas turbine engine section as set forth in  claim 24 , wherein said upstream compressor rotor having an upstream number of stages and said downstream compressor rotor having a downstream number of stages with a ratio of said downstream number of stages to said upstream number of stages being greater than or equal to about 1.6 and less than or equal to about 3.4. 
     
     
         26 . The gas turbine engine section as set forth in  claim 25 , wherein a compression ratio is defined across said upstream and downstream compressor rotors with said compression ratio being greater than or equal to about 38. 
     
     
         27 . The gas turbine engine as set forth in  claim 24 , wherein the fan rotor has less than or equal to 26 blades. 
     
     
         28 . (canceled) 
     
     
         29 . The gas turbine engine as set forth in  claim 27 , wherein a gear ratio of said gear reduction being greater than or equal to about 2.6. 
     
     
         30 . The gas turbine engine as set forth in  claim 29 , wherein the gear ratio of said gear reduction being less than or equal to about 3.3.