Gas turbine engine compressor arrangement

Abstract

A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor, with the overall pressure ratio being greater than or equal to about 35. A pressure ratio across the fan section is less than or equal to about 1.50. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct.

We claim: 
     
         1 . A gas turbine engine comprising:
 a fan section including a fan;   a gear arrangement configured to drive the fan section;   a compressor section, including both a first compressor and a second compressor;   a turbine section configured to drive the compressor section and the gear arrangement;   wherein an overall pressure ratio is:
 provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor; and 
 greater than or equal to about 35; 
   wherein the pressure ratio across the first compressor is greater than or equal to about 7;   wherein a pressure ratio across the fan section is less than or equal to about 1.50; and   wherein the fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct.   
     
     
         2 . The gas turbine engine of  claim 1 , wherein the first compressor is upstream of the second compressor. 
     
     
         3 . The gas turbine engine of  claim 1 , wherein the first compressor is downstream of the second compressor. 
     
     
         4 . The gas turbine engine of  claim 1 , wherein the pressure ratio across the fan section is less than or equal to about 1.45. 
     
     
         5 . The gas turbine engine of  claim 1 , wherein the overall pressure ratio is above or equal to about 50. 
     
     
         6 . The gas turbine engine of  claim 1 , wherein the geared arrangement defines a gear reduction ratio greater than or equal to about 2.3. 
     
     
         7 . The gas turbine engine of  claim 1 , wherein a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, being greater than or equal to about 8. 
     
     
         8 . The gas turbine engine of  claim 1 , wherein the turbine section includes a fan drive turbine configured to drive the fan section, a pressure ratio across the fan drive turbine being greater than or equal to about 5. 
     
     
         9 . The gas turbine engine of  claim 1 , wherein the fan section includes a plurality of fan blades and a fan blade tip speed of each of the fan blades is less than about 1150 ft/second. 
     
     
         10 . A gas turbine engine comprising:
 a fan section including a fan;   a gear arrangement configured to drive the fan section;   a compressor section, including both a first compressor and a second compressor;   a turbine section configured to drive the compressor section and the gear arrangement;   wherein an overall pressure ratio is:
 provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor; and 
 greater than or equal to about 35; 
   wherein the pressure ratio across the first compressor is less than or equal to about 8;   wherein a pressure ratio across the fan section is less than or equal to about 1.50; and   wherein the fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct.   
     
     
         11 . The gas turbine engine of  claim 10 , wherein the first compressor is upstream of the second compressor. 
     
     
         12 . The gas turbine engine of  claim 10 , wherein the first compressor is downstream of the second compressor. 
     
     
         13 . The gas turbine engine of  claim 10 , wherein the pressure ratio across the fan section is less than or equal to about 1.45. 
     
     
         14 . The gas turbine engine of  claim 10 , wherein the overall pressure ratio is above or equal to about 50. 
     
     
         15 . The gas turbine engine of  claim 10 , wherein the pressure ratio across the second compressor is greater than or equal to about 7. 
     
     
         16 . The gas turbine engine of  claim 15 , wherein the pressure ratio across the first compressor is between about 3 and about 8, and the pressure ratio across the second compressor is between about 7 and about 15. 
     
     
         17 . The gas turbine engine of  claim 10 , wherein the geared arrangement defines a gear reduction ratio greater than or equal to about 2.3. 
     
     
         18 . The gas turbine engine of  claim 10 , wherein a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, being greater than or equal to about 8. 
     
     
         19 . The gas turbine engine of  claim 10 , wherein the turbine section includes a fan drive turbine configured to drive the fan section, a pressure ratio across the fan drive turbine being greater than or equal to about 5. 
     
     
         20 . The gas turbine engine of  claim 10 , wherein the fan section includes a plurality of fan blades and a fan blade tip speed of each of the fan blades is less than about 1150 ft/second. 
     
     
         21 . A gas turbine engine comprising:
 a fan section including a fan;   a gear arrangement configured to drive the fan section;   a compressor section, including both a first compressor and a second compressor;   a turbine section configured to drive the compressor section and the gear arrangement;   wherein an overall pressure ratio is provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, the pressure ratio across the first compressor being less than about 8, and the pressure ratio across the second compressor being greater than or equal to about 7;   wherein a pressure ratio across the fan section is less than or equal to about 1.50; and   wherein the fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct.   
     
     
         22 . The gas turbine engine of  claim 21 , wherein the first compressor is upstream of the second compressor. 
     
     
         23 . The gas turbine engine of  claim 21 , wherein the first compressor is downstream of the second compressor. 
     
     
         24 . The gas turbine engine of  claim 21 , wherein the pressure ratio across the fan section is less than or equal to about 1.45. 
     
     
         25 . The gas turbine engine of  claim 21 , wherein the overall pressure ratio is greater than or equal to about 35. 
     
     
         26 . The gas turbine engine of  claim 21 , wherein the geared arrangement defines a gear reduction ratio greater than or equal to about 2.3. 
     
     
         27 . The gas turbine engine of  claim 21 , wherein a bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, being greater than or equal to about 8. 
     
     
         28 . The gas turbine engine of  claim 21 , wherein the turbine section includes a fan drive turbine configured to drive the fan section, a pressure ratio across the fan drive turbine being greater than or equal to about 5. 
     
     
         29 . An arrangement for a gas turbine engine comprising:
 a fan section having a central axis;   a compressor case for housing a compressor;   an inlet case for guiding air to the compressor, the compressor case positioned axially further from the fan section than the inlet case;   a support member extending between the fan section and the compressor case wherein the support member restricts movement of the compressor case relative to the inlet case; and   the compressor case includes a front compressor case portion and a rear compressor case portion, the rear compressor case portion being axially further from the inlet case than the front compressor case portion, wherein the support member extends between the fan section and the front compressor case portion, and the inlet case is removable from the gas turbofan engine separately from the compressor case, the compressor case includes a first compressor section and a second compressor section, and wherein a turbine section drives at least one of the first and second compressor sections, and a gear arrangement is driven by the turbine section such that the gear arrangement drives the fan section, and there being a plumbing connection area positioned upstream of the support member to be utilized for maintenance and repair.