Propulsion system arrangement for turbofan gas turbine engine

Abstract

A method of designing an engine according to an exemplary aspect of the present disclosure includes, among other things, designing an engine and a nacelle assembly together in an interactive process, the engine including a turbine section that drives a fan section and a compressor section. The step of designing the compressor section includes the step of designing a first compressor and a second compressor, with an overall pressure ratio being greater than or equal to about 35. The step of designing the nacelle assembly includes the step of designing the fan section to include a fan nacelle arranged at least partially about a fan, with the fan section having a fan pressure ratio of less than about 1.7. The step of designing the fan section includes configuring the fan section to deliver a portion of air into the compressor section, and a portion of air into a bypass duct, and with a bypass ratio equal to or greater than about 5.

What is claimed is: 
     
         1 . A method of designing an engine, comprising the steps of:
 designing an engine and a nacelle assembly together in an interactive process, the engine including a turbine section that drives a fan section and a compressor section;   wherein designing the compressor section includes the step of designing a first compressor and a second compressor, with an overall pressure ratio provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, the overall pressure ratio being greater than or equal to about 35;   wherein the step of designing the nacelle assembly includes the step of designing the fan section to include a fan nacelle arranged at least partially about a fan, and an aft nacelle moveable relative to a core cowling, with the fan section having a fan pressure ratio of less than about 1.7;   wherein the step of designing the fan section includes arranging the fan section to deliver a portion of air into the compressor section, and a portion of air into a bypass duct defining a bypass flow path, and with 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, equal to or greater than about 5; and   wherein the step of designing the nacelle assembly includes arranging the aft nacelle at least partially about the core cowling to define an aft flowpath portion of the bypass flow path, the aft flowpath portion extending circumferentially between opposite sides of a single bifurcation positioned in the bypass duct such that the aft flowpath portion is uninterrupted by another bifurcation between the opposite sides.   
     
     
         2 . The method as recited in  claim 1 , further comprising designing a mounting assembly attachable to the nacelle assembly and the engine in the interactive process, the mounting assembly including a forward mount, an intermediate mount and an aft mount, wherein:
 the forward mount that reacts to at least a vertical load, a first side load and a torsional load relative to an axis,   the intermediate mount that reacts to at least a thrust load along the axis, and   the aft mount that reacts to at least a vertical load and a second side load.   
     
     
         3 . The method as recited in  claim 2 , wherein the nacelle assembly includes a hardwall containment system. 
     
     
         4 . The method as recited in  claim 3 , wherein the nacelle assembly includes a fan case made of an organic matrix composite. 
     
     
         5 . The method as recited in  claim 4 , wherein the turbine section includes a fan drive turbine that drives the fan section, and a pressure ratio across the fan drive turbine is greater than or equal to about 5. 
     
     
         6 . The method as recited in  claim 5 , wherein the nacelle assembly includes a noise attenuating nozzle having at least one serration defining a trailing edge. 
     
     
         7 . The method as recited in  claim 1 , wherein the nacelle assembly a slim-line nacelle. 
     
     
         8 . The method as recited in  claim 7 , wherein the slim-line nacelle defines a maximum diameter and has an inlet lip defining a highlight diameter, and a ratio of the highlight diameter to the maximum diameter is greater than or equal to about 0.80. 
     
     
         9 . The method as recited in  claim 8 , wherein the aft mount is attachable to a mid-turbine frame of the engine. 
     
     
         10 . The method as recited in  claim 9 , wherein the nacelle assembly includes a noise attenuating nozzle having at least one serration defining a trailing edge. 
     
     
         11 . The method as recited in  claim 10 , wherein the ratio of the highlight diameter to the maximum diameter is greater than or equal to about 0.90. 
     
     
         12 . The method as recited in  claim 1 , wherein the turbine section includes a fan drive turbine that drives the fan section, and a pressure ratio across the fan drive turbine is greater than or equal to about 5. 
     
     
         13 . The method as recited in  claim 12 , wherein the step of designing the nacelle assembly includes the step of designing a thrust reverser to selectively communicate air from the bypass duct. 
     
     
         14 . The method as recited in  claim 13 , further comprising designing a mounting assembly attachable to the nacelle assembly and the engine in the interactive process, the mounting assembly including a forward mount, an intermediate mount and an aft mount, wherein:
 the forward mount that reacts to at least a vertical load, a first side load and a torsional load relative to an axis,   the intermediate mount that reacts to at least a thrust load along the axis, and   the aft mount that reacts to at least a vertical load and a second side load.   
     
     
         15 . The method as recited in  claim 14 , wherein the step of designing the mounting assembly includes attaching the forward mount to the fan nacelle. 
     
     
         16 . The method as recited in  claim 15 , wherein only the forward mount relative to the mounting assembly reacts to torsional loads. 
     
     
         17 . The method as recited in  claim 14 , wherein the intermediate mount is attachable to an engine intermediate case. 
     
     
         18 . The method as recited in  claim 17 , wherein the step of designing the nacelle assembly includes designing a plurality of circumferentially spaced struts to extend in the bypass flow path radially between the engine intermediate case and a fan case surrounding the fan, and positioning the plurality of struts axially between the fan and the single bifurcation. 
     
     
         19 . The method as recited in  claim 14 , wherein the step of designing the mounting assembly includes attaching the aft mount to a mid-turbine frame. 
     
     
         20 . The method as recited in  claim 14 , wherein the step of designing the mounting assembly includes attaching the aft mount to a turbine exhaust case.