Patent Publication Number: US-11391216-B2

Title: Elongated geared turbofan with high bypass ratio

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 14/038,886, filed Sep. 27, 2013, which is a divisional of U.S. patent application Ser. No. 13/792,303, filed Mar. 11, 2013, which claims priority to U.S. Provisional Patent Application Ser. No. 61/761,359 filed Feb. 6, 2013. 
    
    
     BACKGROUND 
     Gas turbine engines are known, and when utilized on an airplane, typically have a fan that delivers air both into a bypass duct defined inwardly of a nacelle and into a core duct the leads to a compressor. The air is compressed in the compressor and delivered into a combustor where it is mixed with fuel and ignited. Products of this combustor pass downstream over turbine rotors driving them to rotate. The turbine rotors, in turn, drive the fan and the compressor. 
     Historically, a fan drive turbine drove both the fan and a low pressure compressor through a direct drive connection such that all of the fan drive turbine, the fan, and the low pressure compressor rotated at the same angular velocity. By tying the speed of the fan to the fan drive turbine, this not only limited the speed of the fan drive turbine, but also was a design restriction on the diameter and speed of the fan. For many reasons, it would be desirable for the fan to rotate at a slower speed, thereby enabling it to be radially larger. 
     More recently, a gear reduction has been provided between the fan drive turbine and the fan. The gear reduction has allowed the fan diameter to increase dramatically. With the increase in fan&#39;s diameter, a bypass ratio, or volume of air delivered into the bypass duct compared to the volume of air delivered into the core duct that leads to the compressor, has also increased. As a result of the increase in the bypass ratio, negative aerodynamic effects have been identified in the overall propulsion system that includes the nacelle and the engine. Accordingly, what is needed is an improved propulsion system that does not experience these negative aerodynamic effects. 
     SUMMARY 
     In a featured embodiment a propulsion system has a fan and a gear. A turbine is configured to drive the gear to drive the fan. The turbine has an exit point. A diameter (D t ) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance (L c  or L n ) from the exit point. A ratio of the distance (L c  or L n ) to the diameter (D t ) is greater than or equal to about 0.90. 
     In another embodiment according to the previous embodiment, the core engine exhaust nozzle includes a plug. The downstream most point of the core engine exhaust nozzle is defined by a downstream end of the plug. The ratio is greater than or equal to about 1.06. 
     In another embodiment according to any of the previous embodiments, the ratio is greater than or equal to about 1.20. 
     In another embodiment according to any of the previous embodiments, a plug is received within the core engine exhaust nozzle. A downstream end of the core engine exhaust nozzle extends downstream of a downstream most end of the plug. The distance (L n ) is defined to a downstream most end of the core engine exhaust nozzle. The ratio is greater than or equal to about 1.02. 
     In another embodiment according to any of the previous embodiments, the ratio is greater than or equal to about 1.17. 
     In another embodiment according to any of the previous embodiments, a bypass ratio is greater than about 6. 
     In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 10. 
     In another embodiment according to any of the previous embodiments, an exhaust case is positioned between the turbine and the core engine exhaust nozzle. 
     In another featured embodiment, a propulsion system has a fan and a gear. A turbine is configured to drive the gear to drive the fan. The turbine has an exit point. A diameter (D t ) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. The core engine exhaust nozzle has a plug. A downstream most point of the core engine nozzle is defined by a downstream end of the plug at a distance (L c ) from the exit point. A ratio of the distance (L c ) to the diameter (D t ) is greater than or equal to about 1.06. 
     In another embodiment according to the previous embodiment, the ratio is greater than or equal to about 1.20. 
     In another embodiment according to any of the previous embodiments, an exhaust case is positioned between the exit of the turbine and an entrance to the engine exhaust nozzle. 
     In another embodiment according to any of the previous embodiments, a bypass ratio is greater than about 6. 
     In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 10. 
     In another featured embodiment, a propulsion system has a fan and a gear. A turbine is configured to drive the gear to drive the fan. The turbine has an exit point. A diameter (D t ) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is downstream of the exit point. A downstream most point of the core engine exhaust nozzle is downstream of an internal plug received within the core engine exhaust nozzle. The downstream most point is defined at a distance (L n ) from the exit point. A ratio of the distance (L n ) to the diameter (D t ) is greater than or equal to about 0.90. 
     In another embodiment according to the previous embodiment, the ratio is greater than or equal to about 1.02. 
     In another embodiment according to any of the previous embodiments, the ratio is greater than or equal to about 1.17. 
     In another embodiment according to any of the previous embodiments, an exhaust case is positioned between the exit of the turbine and an entrance to the engine exhaust nozzle. 
     In another embodiment according to any of the previous embodiments, a bypass ratio is greater than about 6. 
     In another embodiment according to any of the previous embodiments, the bypass ratio is greater than about 10. 
     In another embodiment according to any of the previous embodiments, a gear ratio of the gear is greater than or equal to about 2.3. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  shows a first embodiment of an improved propulsion system according to the present invention. 
         FIG. 3  shows a second embodiment of an improved propulsion system according to the present invention. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  of  FIG. 1  is 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  31  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 low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in the 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 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged in the 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  50  may be varied. For example, gear architecture  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 six (6:1), with an example embodiment being greater than about ten (10:1), 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 (2.3:1) and the low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). 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. 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. 
     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 1 bf 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. “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 high bypass ratio engines, a nacelle  102  as shown in  FIG. 2  will have a relatively great diameter. A fan  100  is shown within the nacelle  102 , somewhat schematically. The fan is driven by a gear drive  112 , and driven by a fan drive turbine  146  in a turbine section  116 . Turbine section  116  may include a higher pressure turbine  147  upstream of the fan drive turbine  146 . A compressor  114  is also illustrated. A diameter D t  is defined as the diameter of the last blade airfoil stage  117  in the fan drive turbine section  146 . 
     A core engine exhaust nozzle  122  has an inner periphery  124  which tapers downwardly to define a nozzle at an end point  125 . The angle at which the nozzle tapers has a maximum defined by balancing aerodynamic characteristics and core engine exhaust nozzle weight. As an example, the maximum angle may be approximately greater than twelve degrees or less than seventeen degrees, and preferably between fourteen and sixteen degrees, and most preferably at fifteen degrees, all measured relative to the horizontal. 
     A plug  126  is shown to extend beyond an end point  125  of a housing of the core engine exhaust nozzle  122 . The plug has a downstream most end  128 . 
     The use of a gear drive  112  reduces the overall length of the turbine section  116  as compared to conventional direct drive turbofan engines. As an example, a direct drive turbofan engine capable of producing a similar amount of thrust as the engine embodiment shown in  FIG. 2 , may have its last turbine airfoil stage at the point  120  (schematically shown). Further, such a conventional direct drive turbofan engine typically would have a nacelle  110  (schematically shown) with a much smaller diameter as compared to the nacelle  102  of the engine embodiment shown in  FIG. 2 . 
     The nacelle  102  has a maximum diameter at point  104 . To eliminate (or at least reduce) negative aerodynamic effects, an outer surface  106  of the nacelle  102 , which is downstream of the point  104 , also has a limitation on a maximum inwardly extending angle to prevent separation of air, balancing aerodynamic characteristics and nacelle weight. Thus, in one embodiment, the maximum angle for the surface  106  may be on the order of about fourteen degrees, again measured relative to a horizontal axis. Of course, in other embodiments, the angle may be less than fourteen degrees. 
     An inner surface  108  of the nacelle  102  forms a nozzle at its downstream end  109  with an outer surface  111  of a core housing. In accordance with, conventional gas turbine design principles, manufacturers would typically try to reduce weight, and thus increase fuel efficiency. Under such conventional design strategy, one of ordinary skill would typically seek to minimize the length of the core engine exhaust nozzle  122  and any exhaust case  118 . That is, one might seek to minimize the length downstream of the downstream end  117  of the turbine section  116  illustrated in  FIG. 2 . However, Applicant has discovered that given the maximum angle for the surface  124 , this would raise challenges with regard to creating an effective nozzle at point  109 . To overcome this detriment, the shown embodiment increases the length of the combined exhaust case  118  and core engine exhaust nozzle  122 . While the core engine exhaust nozzle  122  is illustrated starting at the point  120  at which the last turbine airfoil stage of a non-geared engine would be expected to be, this is merely for illustration simplicity. The two points need not be related. The same is true with the illustration that point  120  coincides with the downstream end of an exhaust case  118 . As in clear, the exhaust case  118  expands radially outwardly to point  120  while the exhaust nozzle  122  extends radially inwardly from point  120 . 
     As a result, whereas the overall length of the turbine section  116  of the embodiment shown in  FIG. 2  is shorter than the corresponding length of the turbine of a non-geared counterpart engine, the overall length of the combined exhaust case  118  and nozzle  122  of the embodiment shown in  FIG. 2  is longer than would be expected. 
     To define the length of the nozzle  122  and exhaust case  118  (if used), a dimension Lc is defined from the point  117  to the point  128 . 
     As an example, in one engine, D t  was 27.6 in., and L c  was 33.5 in. This results in a ratio of about 1.21. In another engine example, where D t  was 33.5 in. and L c  was 43.7 in., the ratio was about 1.30. In a third engine example, where D t  was 35.9 in. and Lc was 50.0 in., the ratio was about 1.39. In another proposed engine example, where D t  was 53.6 in. and L c  was 88.0 in., the ratio was as high as about 1.64. 
     In general, this disclosure extends to geared turbofan engines with a ratio of L c  to D t  of equal to or above about 1.06, and more narrowly equal to or above about 1.20. 
       FIG. 3  shows another embodiment, which is generally the same as the  FIG. 2  embodiment, other than the plug  226  does not extend beyond the downstream end  225  of a housing of the core engine exhaust nozzle  222 . Again, in the shown embodiment, the inward movement of the surface  224  in the nozzle is limited to a maximum angle of about fifteen degrees measured relative to the horizontal, and thus an exhaust case  118  is also utilized in this embodiment. A dimension L n  is defined between the point  117  at the downstream end of the fan drive turbine section  146  of the turbine section  116  and the point  225  at the downstream end of the nozzle  222 . 
     In one such engine example, where D t  was 27.6 in. and L n  was 28.2 in., the ratio was about 1.02. In another engine example, where D t  was 33.5 in. and L n  was 34.6 in., the ratio was about 1.03. In another engine example, wherein D t  was 35.9 in. and L n  was 38.8 in., the ratio was about 1.08. In another proposed engine, where D t  was 53.6 in. and L n  was 69.2 in., the ratio was about 1.29. 
     In general, this disclosure extends to geared turbofan engines with a ratio of Ln to D t  equal to or above about 0.90, more narrowly above about 1.02, and more narrowly above about 1.17. 
     For purposes of this application, the plug and housing are collectively part of a core engine exhaust nozzle, such that points  128  and  225  are the respective downstream most points of the core engine exhaust nozzle. 
     The core engine exhaust nozzle itself should have sufficient stiffness, and should be formed of a material that would have appropriate strength characteristics at 1,200° F. A material with a density of about 0.3 lbs./in. 3  may be utilized to reduce the overall weight. In one embodiment, the core engine exhaust nozzle  122 / 222  may be formed of rolled sheet stock, with a thickness less than 2.5 percent of a diameter of an inner flow path of a turbine. In another embodiment, the core nozzle may be formed of a sandwich structure, or may be formed to have a corrugated shape to reduce weight. In another embodiment, the core engine exhaust nozzle may be formed of ceramic matrix composites. Of course, other materials for the core exhaust nozzle are possible and are fully within the scope of this disclosure. 
     Although various embodiments of this invention have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.