Patent Publication Number: US-11661906-B2

Title: Propulsion system arrangement for turbofan gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 16/254,889, filed on Jan. 23, 2019, which is a continuation of U.S. patent application Ser. No. 15/880,717, filed on Jan. 26, 2018, which is a divisional of U.S. patent application Ser. No. 14/615,484, filed on Feb. 6, 2015 and issued as U.S. Pat. No. 9,915,225. 
    
    
     BACKGROUND 
     The present disclosure relates to a propulsion system having integrated designs for at least a nacelle and a turbofan gas turbine engine. Preferably, such integrated design accounts for an engine, an engine nacelle, and structure for connecting an engine to a wing, such as an engine pylon or portions thereof. 
     The aerospace industry typically approaches the design of a wing, pylon, nacelle and engine in a disjoined, hierarchical or isolated manner. Engine designers must often meet design requirements specified by airframe customers. This approach may constrain or undesirably affect the design or implementation of the engine, nacelle, mounting, and various related structure. 
     Designing an integrated propulsion system requires optimizing force distributions and reactions at engine mounts. A gas turbine engine may be mounted at various points on an aircraft, such as a pylon integrated with an aircraft structure. An engine mounting configuration ensures the transmission of loads between the engine and the aircraft structure. The loads typically include the weight of the engine, its thrust, aerodynamic loads, maneuver loads, and rotary torque about the engine axis. The engine mounting configuration must also absorb the deformations to which the engine is subjected during different flight phases, and for example, the dimensional variations due to thermal expansion and retraction. 
     SUMMARY 
     A method of designing an engine according to an example of the present disclosure includes the steps of designing an engine and a nacelle assembly together in an interactive process. The engine includes a turbine section configured to drive a fan section and a compressor section. 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 is 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, 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. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly includes a single bifurcation in the bypass duct. 
     A further embodiment of any of the foregoing embodiments includes designing a mounting assembly attachable to the nacelle assembly and the engine in the interactive process. The mounting assembly includes a forward mount, an intermediate mount and an aft mount. The forward mount is configured to react to at least a vertical load, a first side load and a torsional load relative to an axis. The intermediate mount is configured to react to at least a thrust load along the axis. The aft mount is configured to react to at least a vertical load and a second side load. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly includes a hardwall containment system. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly includes a fan case made of an organic matrix composite. 
     In a further embodiment of any of the foregoing embodiments, the turbine section includes a fan drive turbine configured to drive the fan section. A pressure ratio across the fan drive turbine is greater than or equal to about 5. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly includes a noise attenuating nozzle having at least one serration defining a trailing edge. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly is a slim-line nacelle. 
     In a further embodiment of any of the foregoing embodiments, the slim-line nacelle defines a maximum diameter and has an inlet lip defining a highlight diameter. A ratio of the highlight diameter to the maximum diameter is greater than or equal to about 0.80. 
     In a further embodiment of any of the foregoing embodiments, the aft mount is attachable to a mid-turbine frame of the engine. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly includes a noise attenuating nozzle having at least one serration defining a trailing edge. 
     In a further embodiment of any of the foregoing embodiments, the ratio of the highlight diameter to the maximum diameter is greater than or equal to about 0.90. 
     An integrated propulsion system according to an example of the present disclosure includes components that include a gas turbine engine and a nacelle assembly. The system designed by a process includes identifying two or more of internal structural loading requirements, external structural mount loading requirements, aerodynamic requirements, and acoustic requirements for the system, and interdependently designing the components to meet the requirements. The nacelle assembly includes a fan nacelle arranged at least partially about a fan and an engine. The system includes a fan section at least partially defined by the fan nacelle, and the fan section has a fan pressure ratio of less than about 1.7. The engine includes a turbine section configured to drive the fan section and a compressor section. The compressor section includes a first compressor and a second compressor. 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 the overall pressure ratio is greater than or equal to about 35. The fan section is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio, defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is equal to or greater than about 5. 
     A further embodiment of any of the foregoing embodiments includes a pylon attaching the nacelle assembly and the engine to an aircraft structure. 
     A further embodiment of any of the foregoing embodiments includes a thrust reverser configured to selectively communicate air from the bypass duct. 
     In a further embodiment of any of the foregoing embodiments, the nacelle assembly is a slim-line nacelle. The slim-line nacelle defines a maximum diameter and has an inlet lip defining a highlight diameter. A ratio of the highlight diameter to the maximum diameter is greater than or equal to about 0.80. 
     In a further embodiment of any of the foregoing embodiments, a single bifurcation is positioned in the bypass duct. 
     In a further embodiment of any of the foregoing embodiments, the pylon includes a mounting assembly attachable to the nacelle assembly and the engine. The mounting assembly includes a forward mount, an intermediate mount and an aft mount. The forward mount is configured to react to at least a vertical load, a first side load and a torsional load relative to an axis. The intermediate mount is configured to react to at least a thrust load along the axis. The aft mount is configured to react to at least a vertical load and a second side load. 
     In a further embodiment of any of the foregoing embodiments, the intermediate mount is attachable to an engine intermediate case. 
     In a further embodiment of any of the foregoing embodiments, the aft mount is attachable to one of a mid-turbine frame and a turbine exhaust case. 
     In a further embodiment of any of the foregoing embodiments, the slim-line nacelle includes a fan nacelle arranged at least partially about a core cowling to define a bypass flow path. An aft nacelle is arranged at least partially about the core cowling to define a duct. The bypass flow path extends circumferentially about the axis between opposite sides of a single bifurcation defining the duct. 
     In a further embodiment of any of the foregoing embodiments, the forward mount is attachable to the fan nacelle. 
     In a further embodiment of any of the foregoing embodiments, the aft mount is attachable to the mid-turbine frame. 
     An integrated propulsion system for a gas turbine engine according to an example of the present disclosure includes a nacelle assembly and a mounting assembly attachable to the nacelle assembly. The mounting assembly includes a forward mount, an intermediate mount and an aft mount. The forward mount is configured to react to at least a vertical load, a first side load and a torsional load relative to an axis. The intermediate mount is configured to react to at least a thrust load along the axis. The aft mount is configured to react to at least a vertical load and a second side load. 
     In a further embodiment of any of the foregoing embodiments, the intermediate mount is attachable to an engine intermediate case. 
     In a further embodiment of any of the foregoing embodiments, the aft mount is attachable to one of a mid-turbine frame and a turbine exhaust case. 
     In a further embodiment of any of the foregoing embodiments, the nacelle includes a core cowling defined about the axis, a fan nacelle arranged at least partially about the core cowling to define a bypass flow path, and an aft nacelle arranged at least partially about the core cowling to define a duct. The bypass flow path extends circumferentially about the axis between opposite sides of a single bifurcation defining the duct. 
     In a further embodiment of any of the foregoing embodiments, the forward mount is attachable to the fan nacelle. 
     A further embodiment of any of the foregoing embodiments includes a thrust reverser positioned axially between the fan nacelle and the aft nacelle. The thrust reverser includes a cascade configured to selectively communicate a portion of fan bypass airflow from the bypass flow path. 
     In a further embodiment of any of the foregoing embodiments, a single bifurcation extends radially between the aft nacelle and the core cowling. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently disclosed embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG.  1    shows a schematic view a gas turbine engine along an engine longitudinal axis; 
         FIG.  2    illustrates a cross section view of a portion of a nacelle; 
         FIG.  3 A  is a side cutaway view of an engine casing for the turbofan engine of  FIG.  1    in a deployed position; 
         FIG.  3 B  is a side cutaway view of the engine casing of  FIG.  3 A  in a stowed position; 
         FIG.  3 C  is a side cutaway view of nacelle assembly arrangements for a gas turbine engine; 
         FIG.  4    is a cross section view of the engine casing of  FIGS.  3 A and  3 B ; 
         FIG.  5 A  is a partial cross section view of a thrust reverser and a variable area nozzle in stowed positions; 
         FIG.  5 B  is a partial cross section view of the thrust reverser of  FIG.  5 A  in the stowed position and the variable area nozzle of  FIG.  5 A  in a deployed position; 
         FIG.  5 C  is a partial cross section view of the thrust reverser and the variable area nozzle of  FIG.  5 A  in deployed positions; 
         FIG.  6 A  is a schematic view of an engine support structure; 
         FIG.  6 B  is a free body diagram illustrating loads reacted by the engine support structure of  FIG.  6 A  according to a first embodiment; 
         FIG.  6 C  is a free body diagram illustrating loads reacted by the engine support structure of  FIG.  6 A  according to a second embodiment; 
         FIG.  7 A  illustrates a prior process for designing a propulsive system; 
         FIG.  7 B  illustrates an interactive process for designing a propulsive system; 
         FIG.  8    is a cross-sectional view of a fan blade according to an embodiment; and 
         FIG.  9    schematically shows the arrangement of the low and high spool of  FIG.  1   , along with the fan drive. 
     
    
    
     DETAILED DESCRIPTION 
     An integrated propulsion system  10  generally includes an engine  20  ( FIG.  1   ), a pylon  108  ( FIGS.  3 A to  4   ), a nacelle assembly  82  ( FIGS.  3 A to  4   ), and an aircraft structure  109  ( FIG.  6 A ) such as an aircraft wing or fuselage. The various components of the integrated propulsion system  10  are designed and implemented in an integrated or holistic manner to achieve a new state. The term “integrated propulsion system” is utilized for the purposes of this disclosure to mean multiple propulsion components, such as a nacelle and a gas turbine engine, and in some instances, also an engine pylon and an aircraft static structure, designed or configured in an interactive process according to commonly defined and interrelated sets of requirements, such that the overall performance of the system is optimized. An explanation of this interactive process, with examples, is provided below, for example, with later reference to  FIGS.  7 A and  7 B . This technique desirably leads to higher overall system performance in terms of size, weight and efficiency. 
       FIG.  1    schematically illustrates a gas turbine engine  20  of the kind which could be part of an integrated propulsion system. The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . In  FIG.  1   , the inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . Such a configuration is commonly referred to as a geared turbofan. Importantly, however, for an integrated propulsion system, the design considerations disclosed herein are not limited to a geared turbofan. Rather, the options for the integrated propulsion system disclosed herein are also applicable to direct drive turbofan, where a fan drive turbine directly drives the fan. 
     The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about five (5), or greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans, single spool configurations, and configurations having more than two spools (i.e., low spool, intermediate spool and high spool). 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. In some embodiments, the fan pressure ratio is equal to or less than about 1.80, equal to or less than about 1.70, equal to or less than about 1.60, equal to or less than about 1.55, or equal to or less than about 1.50. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. In some embodiments, the bypass ratio is greater than or equal to about 5:1 or greater than or equal to about 8:1, between 8:1 and 13:1, or greater than or equal to about 12:1. In other embodiments, the gear reduction ratio is greater than about 2.6. In some embodiments, an overall pressure ratio is provided by the combination of a pressure ratio across a low (or first) compressor  44  and a pressure ratio across the high (or second) second compressor  52  is greater than or equal to about 35 measured at maximum climb conditions, or greater than or equal to about 40. In another embodiment, the overall pressure ratio is above or equal to about 50. For the purposes of this disclosure, the pressure ratio of the low (or first) compressor  44  includes the pressure ratio across a fan root or inner portion of the fan blade  42 . 
     The engine static structure  36  for the illustrated integrated propulsion system  10  generally has sub-structures including a case structure  66  often referred to as the engine “backbone.” In some embodiments, the case structure  66  at least partially houses the engine sections  22 ,  24 ,  26 ,  28  and, where applicable, the geared architecture  48 . In an embodiment of the integrated propulsion system  10 , the case structure  66  includes a fan case  68 , an intermediate case (IMC)  70 , a high pressure compressor case  72 , a thrust case  74 , a low pressure turbine case  76 , and a turbine exhaust case  78 . The fan blades  42  are surrounded by the fan case  68 . 
     The core engine case structure  70 ,  72 ,  74 ,  76 ,  78  for the illustrated integrated propulsion system  10  is secured to the fan case  68  at the IMC  70 . The IMC includes a multiple of circumferentially spaced radially extending struts  80 , which radially span the core engine case structure and the fan case  68 . The gas turbine engine  20  is configured to be suspended from an engine pylon  108  (shown in  FIGS.  3 A,  3 B and  4   ). It should be understood that various engines with various case and frame structures may benefit from the teachings of this disclosure. 
     According to an aspect of the disclosed integrated propulsion system  10 , the engine support structure  106  ( FIG.  6 A ) of the pylon  108  is designed to counteract the various loads and forces observed during aircraft operation. Such a design considers and balances the competing requirements of the engine  20 , nacelle assembly  82 , and aircraft structure  109 , as will become apparent from the teachings of this disclosure. 
     Referring to  FIGS.  3 A,  3 B and  4   , a nacelle assembly or engine casing  82  of the integrated propulsion system  10  includes a core nacelle or cowling  84 , a fan nacelle  86 , a duct  88  and a bifurcation  90 . As will be discussed below, the nacelle assembly or engine casing  82  is designed based in part on the various competing operating and design requirements of the engine  20 , pylon  108 , and aircraft structure  109  to improve the overall efficiency and performance of the system. 
     The core cowling  84  for the integrated propulsion system  10  extends circumferentially around and at least partially houses the engine sections  24 ,  26  and  28  and, where applicable, the cowling  84  extends circumferentially around and houses the geared architecture  48 . The core cowling  84  extends axially along the longitudinal axis A between an inlet  92  of the core flowpath (“core inlet”) and a nozzle  94  of the core flowpath (“core nozzle”) downstream of the inlet  92 . 
     The fan nacelle  86  for the integrated propulsion system  10  extends circumferentially around and houses the fan  42  and at least a portion of the core cowling  84 , thereby defining the bypass flowpath B. The fan nacelle  86  extends axially along the longitudinal axis A between an airflow inlet  96  and a bypass nozzle  98  of the bypass flowpath B (“bypass nozzle”) downstream of the inlet  96 . In some embodiments, the fan nacelle  86  is a slim-line nacelle. 
     The fan nacelle  86  defines a highlight diameter  87  extending radially at an inlet lip  91  of inlet  96 , and defines a maximum diameter  89  in a radial direction. The highlight diameter is measured at the maximum axial extent of the inlet lip  91  relative to the engine axis A. In some embodiments, a ratio of the highlight diameter  87  to the maximum diameter  89  is greater than or equal to about 0.80, or greater than or equal to about 0.90, although other ratios are contemplated. The slim-line nacelle configuration of the integrated propulsion system  10  reduces aerodynamic drag and other loads exerted on the engine  20 , engine pylon  108  and or mounting arrangements, and the various aircraft structure, for example. 
     Referring to  FIG.  3   , a geometry of the nacelle assembly  82  can selected to affect the overall propulsive efficiency of the integrated propulsive system  10 , and can be utilized for a gas turbine engine, such as engine  20  shown in  FIG.  1   . For example, the nacelle assembly can be configured as a laminar flow nacelle or a slim line nacelle, as discussed below and indicated by nacelle assemblies  82   a ,  82   b  and  82   c , having various nacelle thicknesses or profiles. The nacelle assembly can also be arranged to have a relatively short inlet, further improving aerodynamic performance of the integrated propulsive system  10  as discussed in detail below. 
     The nacelle assembly can be a laminar flow nacelle as illustrated by nacelle assembly  82   a  (shown in dashed lines) or a slim line nacelle as illustrated by nacelle assembly  82   b . The maximum diameter  89   a  of the laminar flow nacelle  82   a  is relatively greater than the maximum diameter  89   b  of the slim line nacelle  82   b , which is defined at a location axially aft of the maximum diameter  89   a . Nacelle assembly  82   c  illustrates a laminar flow nacelle in which the design parameters of the nacelle assembly  82   c  are constrained by one or more predetermined design requirements of a pylon  108  (shown in  FIGS.  3 A and  3 B ), causing the maximum diameter  89   c  to be defined axially forward of maximum diameter  89   a.    
     The nacelle assembly  82  can also be configured to have a relatively short inlet to improve aerodynamic performance. Fan blades  42  establish a diameter between circumferentially outermost edges  43  of the fan blades  42  corresponding leading edge  45 . The fan diameter D 1  is shown as a dimension extending between the edges  43  of two of the fan blades  42  that are parallel to each other and extending in opposite directions away from the central axis A. A length L 1  is established between inlet lip  91  at the central axis A and an intersection of a plane defining the fan diameter D 1 , with the plane being generally perpendicular to the central axis A. 
     The length L 1  and the fan diameter D 1  are selected based in part on the arrangement of other components of the integrated propulsion system  10 . In some embodiments, the dimensional relationship of L 1 /D 1  (e.g., the ratio of L 1 /D 1 ) is between about 0.2 and about 0.45, and in further embodiments is between about 0.25 and about 0.45, or between about 0.30 and about 0.40. The dimensional relationship of L 1 /D 1  is smaller than that found on typical gas turbine engines. Providing a shorter inlet portion length L 1  facilitates reducing the weight of the engine assembly, and also reduces the overall length of the nacelle assembly  82  and reduces external drag. Additionally, having a shorter inlet reduces the bending moment and corresponding load on the engine structure during flight conditions. 
     The nacelle assembly  82  can also be configured to have a relatively compact overall arrangement or envelope. The nacelle assembly  82 , such as fan nacelle  86 , establishes a maximum diameter  89  in a radial direction as previously discussed. The nacelle assembly  82  establishes a length L 2  extending from a forward most portion of inlet lip  91  relative to the engine axis A to an aftmost portion of the nacelle assembly  82 , such as core cowling  84 . In some embodiments, the dimensional relationship of L 2  to maximum diameter  89  (e.g., the ratio of L 2  over maximum diameter  89 ) is between about 1.2 and about 1.5, and in further embodiments is between about 1.25 and about 1.35. 
     Referring back to  FIGS.  3 A and  3 B , in some embodiments of the integrated propulsion system  10 , the fan nacelle  86  includes a stationary forward portion  100  and an aft nacelle  102 . The aft nacelle  102  is moveable relative to the stationary forward portion  100 , and for example, is configured to selectively translate axially along a supporting structure such as a plurality of guides or tracks  104  ( FIG.  3 A ). The tracks  104  may be connected to opposing sides of an engine mounting configuration or support structure  106 , which may be configured as part of, or attached to, an engine pylon  108  that mounts the turbine engine  20  to an aircraft airframe, such as an aircraft wing or fuselage. In some embodiments, the aft nacelle  102  includes a noise attenuating nozzle which may have various noise attenuating materials or may have one or more serrations  93  ( FIG.  3 A ) defining a trailing edge of the bypass nozzle  98 . 
     As seen in  FIG.  4   , the bifurcation  90  extends radially between the core cowling  84  and the fan nacelle  86 . In some embodiments, the bifurcation  90  is circumferentially aligned with and at least partially receives the engine support structure  106 . The bifurcation  90  is positioned radially between the engine axis A and the pylon  108 , and may be referred to as an upper bifurcation, or upper “bifi”. The bifurcation  90  is not limited to any particular spatial orientations, and may alternatively be located within a gravitational side or bottom portion of the duct  88 . The bifurcation  90  extends radially between the core cowling  84  and the fan nacelle  86  through the bypass flowpath B, thereby bifurcating the bypass flowpath B. The bifurcation  90  is aerodynamically contoured to define a portion of the bypass flowpath B, and in some embodiments, defines at least a portion of the support structure  106 . 
     In the integrated propulsion system of  FIG.  4   , only an upper bifi is utilized. Load sharing normally associated with an upper and lower bifi has to be managed with the omission of the lower bifi, as would be appreciated by one of ordinary skill reading this disclosure. Such a configuration, however, reduces aerodynamic drag of using both an upper and lower bifi. Thus, the solution of a single bifi in the illustrated integrated propulsion system  10  derives certain advantages over a multi-bifi system. 
     A radially inner portion of the duct  88  is bounded by at least a portion of the core cowling  84 . A radially outer portion of the duct  88  is bounded by the aft nacelle  102 . The duct  88  is bounded in the circumferential direction by the bifurcation  90  to define a bypass duct portion for the illustrated integrated propulsion system  10 , which defines at least an aft portion C of the bypass flowpath (“aft flowpath portion”). In some embodiments, the aft flowpath portion C is substantially uninterrupted by bifurcation(s) and/or support structure(s) other than the bifurcation  90  and/or the support structure  106 . The flow path circumferentially defines an angle Θ between opposing surfaces  110  and  112  of the bifurcation  90  as measured, for example, at a widest portion of the bifurcation  90  within the duct  88 . In some embodiments, the angle Θ is greater than or equal to about 180 degrees. In other embodiments, the angle Θ is greater than or equal to about 320 degrees, although other circumferential bounds of the aft flowpath portion C are contemplated. 
     The aft flowpath portion C extends axially within the aft nacelle  102  to the bypass nozzle  98  (shown in  FIGS.  3 A and  3 B ). The aft flowpath portion C extends radially between a radial outer surface  114  of the core cowling  84  and a radial inner surface  116  of the aft nacelle  102 . Although a bypass duct portion having a single bifurcation in the circumferential direction, sometimes referred to as an “O-duct,” is disclosed for the illustrated integrated propulsion system  10 , it should be appreciated that other duct arrangements are contemplated, including D-duct arrangements having two bifurcations (e.g., extending less than or equal to about 180 degrees around a longitudinal axis A). 
     Referring to  FIGS.  5 A- 5 C , the engine casing  82  in some embodiments of the integrated propulsion system  10  includes at least one thrust reverser  118  and/or a variable area nozzle (VAN)  120 , sometimes referred to as a variable area fan nozzle (VAFN), for adjusting various characteristics of the bypass flow.  FIG.  5 A  illustrates the thrust reverser  118  and the variable area nozzle  120  in stowed positions.  FIG.  5 B  illustrates the thrust reverser  118  in the stowed position and the variable area nozzle  120  in a deployed position.  FIG.  5 C  illustrates the thrust reverser  118  in a deployed position and the variable area nozzle  120  in stowed position. The thrust reverser  118  includes a thrust reverser body  122 , which is configured with the aft nacelle  102 . In some embodiments, the thrust reverser  118  includes one or more blocker doors  124 , one or more actuators  126 , and/or one or more cascades  128  of turning vanes  130  arranged circumferentially around the longitudinal axis A. 
     The thrust reverser body  122  for the integrated propulsion system  10  may have a generally tubular geometry with an axially extending slot or channel configured to accommodate the support structure  106  (shown in  FIGS.  3 A- 3 B ). The thrust reverser body  122  includes at least one recess  132  that houses the cascades  128  and the actuators  126  when the thrust reverser  118  is in the stowed position. Each blocker door  124  is pivotally connected to the thrust reverser body  122 . The actuators  126  are adapted to axially translate the thrust reverser body  122  between the stowed and deployed positions. As the thrust reverser body  122  translates aftwards, the blocker doors  124  pivot radially inward into the aft flowpath portion C and divert at least some or substantially all of the bypass air as flow F through the cascades  128  to provide the reverse engine thrust. 
     Other thrust reverser configurations are contemplated for the integrated propulsion system  10 . In some embodiments, the cascades  128  are configured to translate axially with a respective thrust reverser body  122 . In other embodiments, the thrust reverser body  122  and/or cascades  128  include one or more circumferential segments that synchronously or independently translate or otherwise move between deployed and stowed positions. 
     Referring back to  FIGS.  3 A and  3 B , in some embodiments of the integrated propulsion system  10 , the thrust reverser  118  is configured without blocker doors. Opposing surfaces  114 ,  116  of the core cowling  84  and/or aft nacelle  102  may include one or more contoured segments to define a radial distance  131 . As the aft nacelle  102  translates aftwards, the radial distance  131  of the aft flowpath portion C may change (e.g., reduces) to a radial distance  131 ′ to partially or fully obstruct the bypass flowpath B to divert flow through the cascade  128  (shown in dashed lines at the bottom of  FIG.  3 A ) to provide the reverse engine thrust. 
     Referring to  FIGS.  5 A to  5 C , the variable area nozzle  120  for the illustrated integrated propulsion system  10  includes a nozzle body  134  and one or more actuators  135 . The nozzle body  134  is configured with the aft nacelle  102 , and is arranged radially within and may nest with the thrust reverser body  122 . The nozzle body  134  may have a generally tubular geometry with an axially extending slot or channel configured to accommodate the support structure  106  (shown in  FIGS.  3 A- 3 B ). The actuators  135  are configured to axially translate the nozzle body  134  between the stowed position of  FIG.  5 A  and the deployed position of  FIG.  5 B . As the nozzle body  134  translates aftwards, a radial distance  137  of the bypass nozzle  98  between a trailing edge or aft end  136  of the fan nacelle  86  and the core cowling  84  may change (e.g., increase) and thereby change (e.g., increase) a flow area of the bypass nozzle  98 . In this manner, the variable area nozzle  120  may adjust a pressure drop or ratio across the bypass flowpath B (see  FIGS.  3 A and  3 B ) by changing the flow area of the bypass nozzle  98 . 
     In some embodiments, the variable area nozzle  120  defines or otherwise includes at least one auxiliary port  138  to affect the bypass flow. In some embodiments, the auxiliary port  138  is defined between an upstream portion  140  of the aft nacelle  102  and the nozzle body  134  of the variable area nozzle  120  as the nozzle body  134  translates axially aftwards. A flow area through the auxiliary port  138  augments the flow area of the bypass nozzle  98 , thereby increasing an effective flow area of the variable area nozzle  120 . The variable area nozzle  120  for the illustrated integrated propulsion system  10  therefore may adjust a pressure drop or ratio across the bypass flowpath B while translating the nozzle body  134  over a relatively smaller axial distance. Other variable area nozzle configurations are contemplated. In some embodiments, the variable area nozzle  120  includes one or more bodies (e.g., flaps) that may move radially (or axially and radially) to change the flow area of the bypass nozzle  98 . 
     As will now be further disclosed, the engine support structure  106  ( FIG.  6 A ) of the pylon  108  for the illustrated integrated propulsion system  10  is designed to counteract the various loads and forces observed during aircraft operation.  FIG.  6 A  illustrates a highly schematic view of an engine support structure  106  according to an embodiment of the integrated propulsion system  10 . The engine support structure  106  includes a forward mount  142 , an intermediate mount  144 , and an aft mount  146 . The forward mount  142  is attachable to at least one of a fan nacelle  86 , another portion of a pylon  108  and an aircraft interface or structure  109 . The intermediate mount  144  is attachable to an intermediate case  70 . 
     In one embodiment, the aft mount  146  is attachable to a turbine exhaust case  78 . In another embodiment, the aft mount  146  is attachable to a mid-turbine frame  57 . Attaching the aft mount  146  to the mid-turbine frame  57  generally increases ground clearance by moving portions of the engine  20  relatively closer to the aircraft wing. In some embodiments, at least one of the intermediate mount  144  and the aft mount  146  is attachable to another portion of the pylon  108  or aircraft structure  109 . In one embodiment, the forward mount  142  is attachable to the intermediate mount  144 , and in another embodiment, the intermediate mount  144  is attachable to the aft mount  146 . 
       FIG.  6 B  illustrates a free body diagram of an engine support structure  106 ′ for the integrated propulsion system  10 , with an aft mount  146  attachable to a turbine exhaust case  78 , and which is configured to react to at least vertical loads V, side loads S, thrust loads T 1 , and torsional loads T 2 . In some embodiments, the forward mount  142  is configured to react to at least one or more vertical loads V, one or more side loads S and torsional loads T 2  relative to the engine axis A. The intermediate mount  144  is configured to react to at least a thrust load T 1  along the engine axis A. The aft mount  146  is configured to react to at least a vertical load V and one or more side loads S. In one example embodiment, only the forward mount  142  is configured to react to torsional loads T 2 . The forward mount  142  is configured to react to torsional loads T 2  which for bypass duct portions having a single bifurcation may be observed relatively forward along the engine axis A as compared to D-duct arrangements. 
       FIG.  6 C  illustrates a free body diagram of an engine support structure  106 ′ for the integrated propulsion system  10 , with an aft mount  146  attachable to a mid-turbine frame  57 . In this arrangement, a forward mount  142  is configured to react to at least one or more vertical loads V and one or more side loads S. An intermediate mount  144  is configured to react at least a thrust load T 1  along the engine axis A. The aft mount  146  is configured to react to at least a vertical load V and one or more side loads S. The aft mount  146  is also configured to react to torsional loads T 2  relative to the engine axis A. In an embodiment, only the aft mount  146  is configured to react to torsional loads T 2 . This technique may be utilized, for example, in combination with engines having relatively high bypass ratios, such as any of the bypass ratios of this disclosure. 
     The engine support structures  106 ,  106 ′ for the integrated propulsion system  10  react to two or more vertical loads, two or more side loads, thrust loads and torsional loads, and are therefore referred to as “determinate systems.” However, it should be appreciated that the engine support structures  106 ,  106 ′ may be configured to react to additional loads, sometimes referred to as indeterminate systems. The engine support structures  106 ,  106 ′ reduce the amount of backbone bending in the engine  20  at various points in the case structure  66 , which may be caused in part by the arrangement of the bypass duct portion with respect to the engine axis A. This reduces blade tip clearance requirements in the compressor section  24  and/or turbine section  28 , for example. 
     As illustrated by  FIGS.  6 A to  6 C  and discussed above, the engine support structures  106 ,  106 ′ are designed holistically as part of the integrated propulsion system  10 . In particular, the selection of locations of the various mount points and which loads to react at the various mount points is determined based in part on the operating and design requirements of the related engine  20  ( FIGS.  1  and  1 A ), nacelle assembly  82  ( FIGS.  3 A to  4   ) and aircraft structure  109 . Loads are therefore reacted more optimally than prior segmented or disjoined mounting techniques. 
     The engine support structures  106 ,  106 ′ in the illustrated integrated propulsion system  10  are arranged in a manner that considers the requirements and optimization of the nacelle assembly  82 , including the bypass duct portion  88  and associated aerodynamic loads and other forces. The resulting configuration reduces overall packaging requirements of the engine and other related external devices. This improves overall aerodynamic efficiency of the nacelle assembly  82 , increasing propulsive efficiencies and reducing fuel burn. 
     For example, as the above disclosed integrated propulsion system  10  that utilizes a single bifi, the external devices can be moved inward, e.g., into the upper bifi, and are therefore arranged more compactly relative to the engine. With such placement and compact design, the impact of their associated loads can be reduced. That is, their loads can be more effectively transferred to, and through, the bifurcation  90  and engine support structures  106 ,  106 ′, and other portions of the pylon  108 . The example integrated propulsion system  10  therefore desirably leads to higher overall system performance in terms of size, weight and efficiency. 
     Although the various embodiments and example engine  20 , nacelle assembly  82  and the engine support structures  106 ,  106 ′ of the pylon  108  are discussed with a geared architecture  48 , it should be appreciated that, as indicated above, the engine support structures  106 ,  106 ′ and other features of this disclosure extends to non-geared or direct drive engines. The engine support structures  106 ,  106 ′ and other features of this disclosure may also be applied to other gas turbine engine architectures, including direct drive turbofans, single spool configurations, and configurations having more than two spools (i.e., low spool, intermediate spool and high spool), and also assemblies configured with various duct arrangements, including bypass duct portions having a single bifurcation and D-duct configurations. 
     The various methods of designing an integrated propulsion system  10 , including a turbofan gas turbine engine  20  and a nacelle assembly  82 , according to bypass ratios equal to or greater than about 5, greater than or equal to about 8, or equal to or greater than about 10, or equal to or greater than about 12, overall pressure ratios greater than or equal to about 35, or equal to or greater than about 50, and fan pressure ratios equal to or less than about 1.70, or equal to or less than about 1.80, are designed together in a holistic and integrated manner to realize various synergistic benefits. Various synergies can be realized by further designing an engine  20 , a nacelle assembly  82 , and a pylon  108  together in an integrated manner. Other synergies can be realized by further designing an engine  20 , a nacelle assembly  82 , a pylon  108  and a thrust reverser  118  together in an interactive design process. 
       FIG.  7 A  illustrates designing a propulsive system in a prior process. As shown in  FIG.  7 A , prior propulsive systems are designed in a hierarchical manner, in which the design requirements of an aircraft structure  109  drive the requirements of an engine  20 , pylon  108  and nacelle assembly  82 . The design requirements of the engine  20  depend on the design requirements of the pylon  108 , and the design requirements of the nacelle assembly  82  depend on the design requirements of both the pylon  108  and the engine  20 . In this manner, the engine  20 , nacelle assembly  82 , pylon  108  and aircraft structure  109  are not designed to improve the overall performance of the propulsion system. 
       FIG.  7 B  illustrates a process for designing an integrated propulsive system  10  in an interactive process. As shown, the design requirements of the engine  20 , nacelle assembly  82 , pylon  108  and aircraft structure  109  are selected in an interactive and interdependent manner to improve the overall performance of the propulsion system  10 , utilizing any of the techniques discussed herein. For example, the interactive process includes identifying two or more of internal structural loading requirements, external structural mount loading requirements, aerodynamic requirements, and acoustic requirements for the system, and interdependently designing the components to meet the requirements. Other design and operating requirements can be utilized in the interactive process, such as fuel consumption, or aircraft type such as short-range, mid-range or long-range models. 
     For example, the design of the pylon  108  can be based on considering the interface(s) with the engine  20 , the aircraft structure  109 , or loads between the engine  20  and the aircraft wing or structure  109 . The design of the engine  20  can be based upon the arrangement of the engine case structure  66 , the definition of hard points or arrangement of external components such as engine controls, plumbing or the like, which in part define or drive nacelle diameter and aerodynamic lines of the nacelle assembly  82 . The design of the engine  20  can also be based upon the location of the mount planes, which depend on mount interfaces, or upon the use of subsystems such as an environmental control system (ECS), generator(s), oil or hydraulic pumps and the like. 
     Similarly, design of the nacelle assembly  82  can be based upon the aerodynamic lines of the nacelle assembly  82 , such as the selection of the nacelle to define a relatively short or long bypass duct, or the selection of the nacelle to have a laminar flow inlet or a slim line nacelle. In addition, the nacelle assembly  82  can be designed based on structural considerations, such as the arrangement or configuration of a thrust reverser or variable area fan nozzle (VAFN), the number of bifurcations used in the nacelle design, and duct type for the bypass duct or mount point interfaces, for example. Furthermore, the nacelle design can be based on a ratio of nacelle inlet to nacelle diameter, utilizing techniques identified in this disclosure, to determine the loads between the nacelle assembly  82  and the engine  20 , such as engine back-bone bending, for example. 
     The various synergies include weight reductions relative to prior propulsion systems designed in a segregated or hierarchical manner. Weight reductions may be achieved by designing at least the gas turbine engine  20  and the nacelle assembly  82  together, or also the pylon  108  and/or aircraft structure  109  together, in an integrated manner, including any of the techniques discussed herein. 
     Weight reduction may be achieved by configuring the nacelle assembly  82  to have a bypass duct portion with a single bifurcation  90 , as discussed above. This technique eliminates one or more bifurcations and supporting structure utilized in prior designs. The single bifurcation  90  is realized in part by designing the pylon  108  together with the nacelle assembly  82 , enabled in part by a mounting assembly, such as engine support structures  106 ,  106 ′, configured to react to various loads applied to the integrated propulsion system  10  as previously discussed. The configuration of the mounting assembly is based in part on the vertical, side, torsional and thrust loads relative to the arrangement of the engine  20  and the nacelle assembly  82 . 
     Configuring the nacelle assembly  82  and pylon  108  in this manner enables the efficient packaging of various engine external components in a more efficient manner, such as an oil pump, thermal management system, environmental control system, or another component driven by an auxiliary gearbox, for example. Packaging of the external engine components may be considered in an integrated manner in combination with the configuration of the mounting assembly including engine support structures  106 ,  106 ′ to efficiently react to the various loads observed during engine operation and provide greater engine stability. The external engine components may be located axially forward or aft of the geared architecture  48  based in part on the arrangement of the mounting assembly, and in turn, the mounting assembly may be configured based on the location of the external engine components in an integrated and holistic manner. 
     Attaching the engine support structure  106  at the mid-turbine frame  57  further simplifies the arrangement of the turbine exhaust case  78 , which may be configured as a non-structural component. Configuring the turbine exhaust case  78  as a non-structural component reduces part counts and the overall weight of the case structure  66  and the integrated propulsion system  10 . 
     Additional weight reduction and reliability can be achieved as compared to prior propulsion systems, by designing the nacelle assembly  82  to include a fan or hardwall containment case  60 . As illustrated in  FIG.  1   , the fan section  22  includes a fan blade or hardwall containment system  60  (shown schematically) arranged about the engine axis A and spaced radially from the fan blades  42 . The hardwall containment system  60  is configured to contain, and absorb the impact of, a fan blade  42  separating from a fan hub  41  or a fragment thereof. In some embodiments, the hardwall containment system  60  is a hard ballistic liner applied to the nacelle or fan case  15 . The hard ballistic liner can include a rigid material such as a resin impregnated fiber structure, metallic structures, or ceramic structures. However, other materials and structures of the hardwall containment system  60  are contemplated. 
     Further weight reductions may be realized by configuring the nacelle assembly  82  and/or hardwall containment case  60  to include various composite materials that may be selected in part based on the arrangement of the integrated propulsion system  10 . As illustrated schematically in  FIG.  2   , the fan case or nacelle  15  is made of an organic matrix composite. The organic matrix composite can include a matrix material  62  and reinforcement fibers  64  distributed through the matrix material  62 . The reinforcement fibers  64  may be discontinuous or continuous, depending upon the desired properties of the organic matrix composite, for example. The matrix material  62  may be a thermoset polymer or a thermoplastic polymer. The reinforcement fibers  64  may include carbon graphite, silica glass, silicon carbide, or ceramic, metal, for example, depending, for example, on rigidity goals. Given this description, one of ordinary skill in the art will recognize that other types of matrix materials and reinforcement fibers may be used. 
     Further weight reductions and propulsive efficiencies may be realized by configuring the fan blades  42  to include lightweight materials. Various lightweight materials can include, but are not limited to, aluminum and composite materials. Various composite materials can include, but are not limited to, two dimensional or three-dimensional composites such as carbon fiber lay-ups or three-dimensional woven carbon fiber and the like. The composite may be formed from a plurality of braided yarns such as carbon fibers. Other materials can be utilized, such as fiberglass, Kevlar®, a ceramic such as Nextel™, and a polyethylene such as Spectra®. The composite can be formed from a plurality of uni-tape plies or a fabric. The fabric can include woven or interlaced fibers, for example. 
     The fan blades  42  may include one or more cores  75 , as shown in  FIG.  8    which is a schematic cross-section of the fan blade in  FIG.  1    through line  8 - 8 . The core  75  can be made from a foam or other lightweight material such as polyurethane. Other materials can be utilized, such as metallic foam and polymethacrylimide (PMI) foam sold under the trade name Rohacell®. The core  75  can be formed from a composite made of one or more plies of fabric or from braided yarns. In alternative embodiments, the fan blades  42  are free of a core. The fan blades  42  can include a sheath  95  located on an exposed surface, such as core  95   a  at a leading edge or core  95   b  at a trailing edge of the fan blade  42 . A sheath  95  can be located at other positions, such as sheath  95   c  along a pressure side P or sheath  95   d  along a suction side S of the fan blade  42 . Various materials of the sheath can be utilized, such as titanium or another material, for example. 
     The overall aerodynamic performance of the integrated propulsion system  10  can be improved relative to prior propulsion systems. Aerodynamic efficiencies may be achieved by designing at least a gas turbine engine  20  and a nacelle assembly  82  together, or also the pylon  108  and/or aircraft structure  109  together, in an integrated manner. 
     Various techniques for designing the integrated propulsion system  10  to improve aerodynamic performance may be utilized. In some examples, the engine  20  and the nacelle assembly  82  are designed together with the aircraft structure  109 , such as a wing or fuselage, to improve the overall aerodynamic interaction of these components. For example, the various axial, radial and/or circumferential locations of the mounts  142 ,  144 ,  146  with respect to the engine axis A may be selected based on the design and operating characteristics of the engine  20  and nacelle assembly  82 , including the various loads to be reacted to at the mounts  142 ,  144 ,  146 , which may be considered together with other aerodynamic design considerations of the integrated propulsion system  10 , such as the desired proximity of the nacelle assembly  82  to the aircraft structure  109 . 
     Aerodynamic efficiencies may be achieved by configuring the nacelle assembly  82  to have a bypass duct portion with a single bifurcation  90 , as discussed above. This technique eliminates one or more bifurcations and supporting structure utilized in prior designs, thereby improving bypass flow characteristics. The single bifurcation  90  is enabled in part by designing the pylon  108  together with the nacelle assembly  82 , such as configuring engine support structures  106 ,  106 ′ according to the various loads applied to the integrated propulsion system  10 , as discussed above. 
     As previously discussed, attaching the engine support structure  106  at the mid-turbine frame  57  simplifies the arrangement of the turbine exhaust case  78 . The turbine exhaust case  78  can be configured to be a non-structural component. In this manner, the turbine exhaust case  78  can be designed to improve the flow characteristics of the bypass flow path and to increase the overall aerodynamic efficiency of the case structure  66  and the integrated propulsion system  10 . 
     The overall aerodynamic efficiency and noise characteristics of the integrated propulsion system  10  can be further improved by configuring the nacelle assembly  82  to include a slim-line nacelle. As previously discussed, the slim-line nacelle defines a maximum diameter and having an inlet lip defining a highlight diameter. A ratio of the highlight diameter to the maximum diameter can be greater than or equal to about 0.65, thereby reducing weight and drag penalties typically associated with nacelles having a “thick” inlet lip. For example, the ratio of the highlight diameter to the maximum diameter can be between about 0.8 and less than 1, where the range floor could alternatively be 0.85 and the range ceiling could be 0.90 or 0.95. For example, the nacelle assembly  82  is configured having a laminar flow nacelle, such that the ratio of highlight diameter to maximum diameter is between about 0.80 and about 0.83. These arrangements provide a relatively efficient aerodynamic profile of the integrated propulsion system  10  by reducing flow separation about the nacelle. The location of the mounting arrangement, including mounts  142 ,  144 ,  146 , can be selected based on the desired aerodynamic characteristics and spatial arrangement of the nacelle. 
     The nacelle assembly  82  can include various passive or active boundary layer control functionality to reduce drag penalty and increase propulsive efficiency. The nacelle assembly  82  can include a transitional nacelle having inlet lip which is selectively moveable or controlled between a first position  91  and a second position  91 ′ (shown schematically in  FIG.  3 A ) during various operational conditions, thereby varying the ratio of highlight diameter to maximum diameter. The geometry of the inlet lip  91 ,  91 ′ can be varied to include any of the ratios or values previously discussed. Various actuation devices  97  can be utilized to vary a geometry of the nacelle assembly  82  during specific flight conditions, including one or more mechanical actuators or a shape memory alloy, for example. 
     The inlet lip can be fixed or passive, having any of the ratios of highlight diameter to maximum diameter previously discussed, to reduce flow separation and related drag penalties. The nacelle assembly  82  can include one or more flow assemblies  99  (shown schematically in  FIG.  3 B ) operable to generate an amount of pressure or suction on exposed surfaces of the nacelle assembly  82  utilizing various techniques. In this manner, the flow assembly  99  reduces flow separation and increases propulsive efficiencies. 
     Other synergies may be achieved by design the various components of the integrated propulsion system  10  based in part on the desired operating characteristics of these components. The gas turbine engine  20  can be configured to further improve the overall efficiency and performance of the integrated propulsion system  10 , such as configuring a fan drive turbine  46  to have pressure ratios greater than or equal to about 5 measured at cruise flight conditions, or more narrowly between about 5 and about 8. The arrangement of other components of the integrated propulsion system  10 , such as the mounting assembly of the pylon  108 , may be designed in part based on the pressure ratios of the fan drive turbine  46  to achieve additional synergies. 
     Other structural, operational and configuration characteristics may be utilized in the iterative design process of achieving the noted pressure ratios. Structurally, an exit area  200  is shown, in  FIGS.  1  and  9   , at the exit location for the high pressure turbine section  54  is the annular area of the last blade of turbine section  54 . An exit area for the low pressure turbine section is defined at exit  201  for the low pressure turbine section is the annular area defined by the last blade of that turbine section  46 . Operationally, as shown in  FIG.  9   , the turbine engine  20  may be counter-rotating. This means that the low pressure turbine section  46  and low pressure compressor section  44  rotate in one direction (“−’), while the high pressure spool  32 , including high pressure turbine section  54  and high pressure compressor section  52  rotate in an opposed direction (“+”). Configurationally, a gear reduction  48  may couple the low pressure turbine section  46  and the fan section  42 . The gear reduction  48 , which may be, for example, an epicyclic transmission (e.g., with a sun, ring, and star gears), is selected such that the fan  42  rotates in the same direction (“+”) as the high spool  32 . With this arrangement, and with the other structure as set forth above, including the various quantities and operational ranges, the stated pressure ratios can be achieved along with very high speeds for the low pressure spool. Under certain circumstances, one may measure low pressure turbine section and high pressure turbine section operation according to a performance quantity that calculates the exit area for the turbine section multiplied by its respective speed squared. This performance quantity (“PQ”) is defined as:
 
 PQ   ltp =( A   lpt   ×V   lpt   2 )  Equation 1:
 
 PQ   hpt =( A   hpt   ×V   hpt   2 )  Equation 2:
 
where A lpt  is the area of the low pressure turbine section at the exit thereof (e.g., at  201 ), where V lpt  is the speed of the low pressure turbine section, where A hpt  is the area of the high pressure turbine section at the exit thereof (e.g., at  200 ), and where V hpt  is the speed of the low pressure turbine section.
 
     Thus, a ratio of the performance quantity for the low pressure turbine section compared to the performance quantify for the high pressure turbine section is:
 
( A   lpt   ×V   lpt   2 )/( A   hpt   ×V   hpt   2 )= PQ   ltp   /PQ   hpt   Equation 3:
 
     PQ lpt /PQ hpt  ratios that provide a turbine section that is smaller than the prior art, both in diameter and axial length, and provide an efficiency of the overall engine that is greatly increased, and in one example can be obtained where areas of the low and high pressure turbine sections are designed as about 558 in 2  and about 91 in 2 , respectively, and speeds of the low and high pressure turbine sections are designed to reach 10179 rpm and 24346 rpm, respectively. 
     That is, having PQ ltp /PQ hpt  ratios in the 0.5 to 1.5 range, or more narrowly, above or equal to about 0.75, 0.8, or 1.0, provide relatively efficient engines. In a direct drive engine, in which the gear reduction  48  is omitted, the ratio can be about 0.2 to about 0.4, or more narrowly between about 0.2 to about 0.3. Although  FIG.  9    illustrates a two spool arrangement, other engine architectures can benefit from the teachings herein, including three spool arrangements. 
     The low pressure compressor section is also improved with this arrangement, and behaves more like a high pressure compressor section than a traditional low pressure compressor section. It is more efficient than the prior art, and can provide more compression in fewer stages. The low pressure compressor section may be made smaller in radius and shorter in length while contributing more toward achieving the overall pressure ratio design target of the engine. 
     Other design characteristics may also be considered in designing the integrated propulsion system  10 , including any of the stage counts and pressure ratios of the compressor section  24  and turbine section  28  discussed herein. The pressure ratios, stage counts and spool speed ranges of the engine  20  may be considered and selected together in an integrated manner with the nacelle assembly  82 , pylon  108  and aircraft structure  109  to improve the overall propulsive efficiency of the integrated propulsion system  10 . 
     Other design considerations for the integrated propulsion system  10  include the torque profile of the various engine  20  components which may be considered in determining the position of each mount  142 ,  144 ,  146  of the support structure  106 . The operating characteristics of the engine  20  components may be selected in part based on the location of the mount  142 ,  144 ,  146  relative to the engine  20  and the nacelle assembly  82 . 
     The overall operating characteristics of the integrated propulsion system  10  can be further improved by designing a portion of the nacelle assembly  82 , such as an aft nacelle  102 . In some examples, the aft nacelle  102  includes a noise attenuating nozzle having at least one serration  93  defining a trailing edge, as previously discussed. In this configuration, the overall noise characteristics of the integrated propulsion system  10  can be further reduced while minimizing weight by reducing the need for noise attenuating materials or structures. In other examples, the noise attenuating nozzle is made of various noise attenuating materials. The materials can include sintered metal, ceramic foam, or a matrix including aramid fibers such as Kevlar®, or another suitable material, for example. The porosity, depth, and material characteristics of the noise attenuating materials may be selected for optimal impedance and thus optimal acoustic attenuation. 
     It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
     The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The disclosed embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.