Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

Abstract

A gas turbine engine assembly includes a fan. A diameter of the fan has a dimension D. The fan has a pressure ratio of greater than 1.20 and less than 1.45. A leading edge on an inlet portion of a nacelle is within a first reference plane oriented at an oblique angle. A forward most portion on the fan blade leading edges is in a second reference plane. A length of the inlet portion has a dimension L different at a plurality of locations on the inlet portion. A geared architecture has a gear reduction ratio of greater than 2.3, a bypass ratio is greater than 10, and a low pressure turbine includes a pressure ratio greater than 5:1. A dimensional relationship of UD is between 0.25 and 0.45. The leading edge on the inlet portion is further from the second reference plane near the top of the assembly.

What is claimed is: 
     
         1 . A gas turbine engine assembly comprising:
 an engine central longitudinal axis;   a fan and a fan case surrounding the fan, the fan including a plurality of fan blades having circumferentially outermost edges, a diameter of the fan having a dimension D extending between the circumferentially outermost edges of the fan blades, each fan blade further having a leading edge, wherein a portion of the fan case is forward of the leading edges of the fan blades, and the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone;   a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, and a leading edge on the inlet portion is within a first reference plane;   a forward most portion on the leading edges of the fan blades is in a second reference plane perpendicular to the engine central longitudinal axis, and the first reference plane is oriented at an oblique angle relative to the second reference plane and to the engine central longitudinal axis, and a length of the inlet portion has a dimension L measured in a direction parallel to the engine central longitudinal axis between a location of the leading edges of the fan blades and the leading edge on the inlet portion;   a compressor section including a low pressure compressor and a high pressure compressor;   a geared architecture with a gear reduction ratio of greater than 2.3;   a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture;   a combustor between the high pressure compressor and the high pressure turbine;   wherein the fan section drives air along a bypass flow path and the compressor section draws air in along a core flow path with a bypass ratio of greater than 10;   wherein the low pressure turbine includes an inlet, an outlet, and a pressure ratio of greater than 5:1, where the pressure ratio is pressure measured prior to the inlet as related pressure measured at the outlet prior to any exhaust nozzle;   wherein a dimensional relationship of L/D is between 0.30 and 0.40; and   wherein the dimension L is different at a plurality of locations on the inlet portion, the leading edge on the inlet portion is further from the second reference plane near the top of the engine assembly than it is near the bottom of the engine assembly, a greatest value of L corresponds to a value of L/D that is at most 0.45; and   a smallest value of L corresponds to a value of L/D that is at least 0.20.   
     
     
         2 . The assembly as recited in  claim 1 , further comprising a low speed spool and a high speed spool, the low speed spool comprising the low pressure compressor and the low pressure turbine, the high speed spool comprising the high pressure compressor and the high pressure turbine, and both the low speed spool and the high speed spool are mounted for rotation about the engine central longitudinal axis relative to a static structure via several bearing systems. 
     
     
         3 . The assembly as recited in  claim 2 , wherein the low speed spool includes an inner shaft that connects the fan and the low pressure turbine. 
     
     
         4 . The assembly as recited in  claim 3 , wherein:
 the inner shaft connects the low pressure compressor and the low pressure turbine; and   the oblique angle is approximately 5 degrees.   
     
     
         5 . The assembly as recited in  claim 3 , wherein the high pressure turbine is a two-stage turbine. 
     
     
         6 . The assembly as recited in  claim 5 , wherein:
 the axial location of the leading edges of the fan blades along the engine central longitudinal axis varies from hub to tip; and   the geared architecture includes an epicyclical gear train that drives the fan at a lower speed than an input speed in the geared architecture.   
     
     
         7 . The assembly as recited in  claim 5 , wherein the fan has a fan tip speed of less than 1150 ft/second. 
     
     
         8 . The assembly as recited in  claim 7 , wherein the turbine section includes a mid-turbine frame between the high pressure turbine and the low pressure turbine, and the mid-turbine frame supports bearing systems in the turbine section. 
     
     
         9 . The assembly as recited in  claim 8 , wherein the mid-turbine frame includes a plurality of vanes in the core flow path. 
     
     
         10 . The assembly as recited in  claim 9 , wherein the plurality of vanes set airflow entering the low pressure turbine during operation. 
     
     
         11 . The assembly as recited in  claim 10 , wherein the plurality of vanes function as inlet guide vanes for the low pressure turbine. 
     
     
         12 . The assembly as recited in  claim 11 , wherein:
 the axial location of the leading edges of the fan blades varies from hub to tip; and   the geared architecture is a planetary gear system.   
     
     
         13 . The assembly as recited in  claim 7 , wherein:
 the axial location of the leading edges of the fan blades varies from hub to tip; and   the oblique angle is approximately 5 degrees.   
     
     
         14 . The assembly as recited in  claim 5 , wherein:
 the low speed spool includes an inner shaft that connects the low pressure compressor and the low pressure turbine; and   the high speed spool includes an outer shaft that connects the high pressure compressor and the high pressure turbine.   
     
     
         15 . The assembly as recited in  claim 14 , wherein:
 the axial location of the leading edges of the fan blades varies from hub to tip;   the geared architecture is a planetary gear system; and   the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone at a cruise design point.   
     
     
         16 . A gas turbine engine assembly comprising:
 an engine central longitudinal axis;   a fan and a fan case surrounding the fan, the fan including a plurality of fan blades having circumferentially outermost edges, a diameter of the fan having a dimension D extending between the circumferentially outermost edges of the fan blades, and each fan blade further having a leading edge, wherein a portion of the fan case is forward of the leading edges of the fan blades, the axial location of the leading edges of the fan blades varies from hub to tip, the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone at a cruise design point, and the fan has a fan tip speed of less than 1150 ft/second;   a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, and a leading edge on the inlet portion is within a first reference plane;   a forward most portion on the leading edges of the fan blades is in a second reference plane perpendicular to the engine central longitudinal axis, the first reference plane is oriented at an oblique angle relative to the second reference plane and to the engine central longitudinal axis, and a length of the inlet portion has a dimension L measured in a direction parallel to the engine central longitudinal axis between a location of the leading edges of the fan blades and the leading edge on the inlet portion;   a compressor section including a low pressure compressor and a high pressure compressor;   a geared architecture including an epicyclical gear train that drives the fan at a lower speed than an input speed in the geared architecture, the geared architecture with a gear reduction ratio of greater than 2.3, and wherein the epicyclical gear train is a planetary gear system;   a turbine section including a low pressure turbine and a two-stage high pressure turbine, the low pressure turbine driving the fan through the geared architecture;   a combustor between the high pressure compressor and the high pressure turbine;   a low speed spool and a high speed spool, the low speed spool comprising the low pressure compressor and the low pressure turbine, the high speed spool comprising the high pressure compressor and the high pressure turbine, both the low speed spool and the high speed spool mounted for rotation about the engine central longitudinal relative to a static structure via several bearing systems, the low speed spool including an inner shaft that connects the fan and the low pressure compressor to the low pressure turbine, and the high speed spool including an outer shaft that connects the high pressure compressor and the high pressure turbine;   wherein the fan section drives air along a bypass flow path and the compressor section draws air in along a core flow path with a bypass ratio of greater than 10;   wherein the low pressure turbine includes an inlet, an outlet, and a pressure ratio of greater than 5:1, where the pressure ratio is pressure measured prior to the inlet as related pressure measured at the outlet prior to any exhaust nozzle;   wherein a dimensional relationship of L/D is between 0.30 and 0.40; and   wherein the dimension L is different at a plurality of locations on the inlet portion, the leading edge on the inlet portion is further from the second reference plane near the top of the engine assembly than it is near the bottom of the engine assembly, a greatest value of L corresponds to a value of L/D that is at most 0.45; and   a smallest value of L corresponds to a value of L/D that is at least 0.20.   
     
     
         17 . The assembly as recited in  claim 16 , wherein the turbine section includes a mid-turbine frame between the high pressure turbine and the low pressure turbine, the mid-turbine frame includes a plurality of vanes in a core flow path, and the plurality of vanes function as inlet guide vanes for the low pressure turbine. 
     
     
         18 . A gas turbine engine assembly comprising:
 an engine central longitudinal axis;   a fan and a fan case surrounding the fan, the fan including a plurality of fan blades having circumferentially outermost edges, a diameter of the fan having a dimension D extending between the circumferentially outermost edges of the fan blades, and each fan blade further having a leading edge, wherein a portion of the fan case is forward of the leading edges of the fan blades, the axial location of the leading edges of the fan blades along the engine central longitudinal axis varies from hub to tip, the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone, and the fan has a fan tip speed of less than 1150 ft/second;   a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, and a leading edge on the inlet portion is within a first reference plane;   a forward most portion on the leading edges of the fan blades is in a second reference plane perpendicular to the engine central longitudinal axis, the first reference plane is oriented at an oblique angle relative to the second reference plane and to the engine central longitudinal axis, and a length of the inlet portion has a dimension L measured in a direction parallel to the engine central longitudinal axis between a location of the leading edges of the fan blades and the leading edge on the inlet portion;   a compressor section including a low pressure compressor and a high pressure compressor;   a geared architecture with a gear reduction ratio of greater than 2.3;   a turbine section including a low pressure turbine and a two-stage high pressure turbine, the low pressure turbine driving the fan through the geared architecture;   a combustor between the high pressure compressor and the high pressure turbine;   wherein the fan section drives air along a bypass flow path and the compressor section draws air in along a core flow path with a bypass ratio of greater than 10;   wherein the low pressure turbine includes an inlet, an outlet, and a pressure ratio of greater than 5:1, where the pressure ratio is pressure measured prior to the inlet as related pressure measured at the outlet prior to any exhaust nozzle;   wherein a dimensional relationship of L/D is between 0.25 and 0.45; and   wherein the dimension L is different at a plurality of locations on the inlet portion, and the leading edge on the inlet portion is further from the second reference plane near the top of the engine assembly than it is near the bottom of the engine assembly.   
     
     
         19 . The assembly as recited in  claim 18 , wherein the geared architecture includes an epicyclical gear train that drives the fan at a lower speed than an input speed in the geared architecture. 
     
     
         20 . The assembly as recited in  claim 19 , wherein the geared architecture is a planetary gear system. 
     
     
         21 . The assembly as recited in  claim 20 , wherein the turbine section includes a mid-turbine frame between the high pressure turbine and the low pressure turbine, and the mid-turbine frame supports bearing systems in the turbine section. 
     
     
         22 . The assembly as recited in  claim 21 , wherein the mid-turbine frame includes a plurality of vanes in a core flow path. 
     
     
         23 . The assembly as recited in  claim 22 , wherein the plurality of vanes set airflow entering the low pressure turbine during operation. 
     
     
         24 . The assembly as recited in  claim 23 , wherein the plurality of vanes function as inlet guide vanes for the low pressure turbine. 
     
     
         25 . The assembly as recited in  claim 22 , wherein the oblique angle is approximately 5 degrees. 
     
     
         26 . The assembly as recited in  claim 25 , wherein the axial location of the leading edges of the fan blades along the engine central longitudinal axis varies from hub to tip. 
     
     
         27 . The assembly as recited in  claim 19 , further comprising a low speed spool and a high speed spool, the low speed spool comprising the low pressure compressor and the low pressure turbine, the high speed spool comprising the high pressure compressor and the high pressure turbine, and both the low speed spool and the high speed spool are mounted for rotation about the engine central longitudinal axis relative to a static structure via several bearing systems. 
     
     
         28 . The assembly as recited in  claim 27 , wherein the low speed spool includes an inner shaft that connects the fan and the low pressure turbine. 
     
     
         29 . The assembly as recited in  claim 28 , wherein the inner shaft connects the low pressure compressor and the low pressure turbine. 
     
     
         30 . The assembly as recited in  claim 29 , wherein the high speed spool includes an outer shaft that connects the high pressure compressor and the high pressure turbine, and the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone at a cruise design point.