Abstract:
A gas turbine engine comprises a gear train defined along an axis. A spool along the axis drives the gear train and includes a low stage count low pressure turbine. A fan i s rotatable at a fan speed about the axis and driven by the low pressure turbine through the gear train. The fan speed is less than a speed of the low pressure turbine. A core is surrounded by a core nacelle defined about the axis. A fan nacelle i s mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A bypass ratio defined by the fan bypass passage airflow divided by airflow through the core is greater than about ten (10).

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present disclosure is a continuation-in-part of U.S. patent application Ser. No. 13/340,988, filed Dec. 30, 2011, which was a continuation-in-part of U.S. patent application Ser. No. 12/131,876, filed Jun. 2, 2008. 
     
    
     BACKGROUND 
       [0002]    The present invention relates to a gas turbine engine and more particularly to an engine mounting configuration for the mounting of a turbofan gas turbine engine to an aircraft pylon. 
         [0003]    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, thrust, aerodynamic side loads, and rotary torque about the engine axis. The engine mount configuration must also absorb the deformations to which the engine is subjected during different flight phases and the dimensional variations due to thermal expansion and retraction. 
         [0004]    One conventional engine mounting configuration includes a pylon having a forward mount and an aft mount with relatively long thrust links which extend forward from the aft mount to the engine intermediate case structure. Although effective, one disadvantage of this conventional type mounting arrangement is the relatively large “punch loads” into the engine cases from the thrust links which react the thrust from the engine and couple the thrust to the pylon. These loads tend to distort the intermediate case and the low pressure compressor (LPC) cases. The distortion may cause the clearances between the static cases and rotating blade tips to increase which may negatively affect engine performance and increase fuel burn. 
       SUMMARY 
       [0005]    In a featured embodiment, a gas turbine engine comprises a gear train defined along an axis. A spool along the axis drives the gear train and includes a low stage count low pressure turbine. A fan i s rotatable at a fan speed about the axis and driven by the low pressure turbine through the gear train. The fan speed is less than a speed of the low pressure turbine. A core is surrounded by a core nacelle defined about the axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A bypass ratio defined by the fan bypass passage airflow divided by airflow through the core is greater than about ten (10). 
         [0006]    In another embodiment according to the previous embodiment, the low stage count includes six or fewer stages. 
         [0007]    In another embodiment according to any of the previous embodiments, the low stage count includes three (3) stages. 
         [0008]    In another embodiment according to any of the previous embodiments, the low stage count includes five (5) stages. 
         [0009]    In another embodiment according to any of the previous embodiments, the low stage count includes six (6) stages. 
         [0010]    In another embodiment according to any of the previous embodiments, the spool is a low spool. 
         [0011]    In another embodiment according to any of the previous embodiments, a fan variable area nozzle is axially movable relative the fan nacelle to vary a fan nozzle exit area and adjust the fan pressure ratio of the fan bypass airflow during engine operation. 
         [0012]    In another embodiment according to any of the previous embodiments, a controller i s operable to control the fan variable area nozzle to vary the fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow. 
         [0013]    In another embodiment according to any of the previous embodiments, the controller is operable to reduce the fan nozzle exit area at a cruise flight condition. 
         [0014]    In another embodiment according to any of the previous embodiments, the controller is operable to control the fan nozzle exit area to reduce a fan instability. 
         [0015]    In another embodiment according to any of the previous embodiments, the fan variable area nozzle defines a trailing edge of the fan nacelle. 
         [0016]    In another embodiment according to any of the previous embodiments, the gear train defines a gear reduction ratio of greater than or equal to about 2.5. 
         [0017]    In another embodiment according to any of the previous embodiments, the gear train defines a gear reduction ratio of greater than or equal to about 2.3. 
         [0018]    In another embodiment according to any of the previous embodiments, the gear train defines a gear reduction ratio of greater than or equal to 2.5. 
         [0019]    In another embodiment according to any of the previous embodiments, the low pressure turbine defines a low pressure turbine pressure ratio that is greater than about five (5). 
         [0020]    In another embodiment according to any of the previous embodiments, the low pressure turbine defines a low pressure turbine pressure ratio that is greater than five (5). 
         [0021]    In another embodiment according to any of the previous embodiments, the low pressure turbine is one of three turbine rotors. The low pressure turbine drives the fan, while the other two of the turbine rotors each drive a compressor section 
         [0022]    In another embodiment according to any of the previous embodiments, a high pressure turbine is also included, with each of the low pressure turbine and the high pressure turbine driving a compressor rotor. 
         [0023]    In another embodiment according to any of the previous embodiments, the gear train is positioned intermediate a compressor rotor driven by the low pressure turbine and the fan. 
         [0024]    In another embodiment according to any of the previous embodiments, the gear train is positioned intermediate the low pressure turbine and the compressor rotor is driven by the low pressure turbine. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0025]    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: 
           [0026]      FIG. 1A  is a general schematic sectional view through a gas turbine engine along the engine longitudinal axis; 
           [0027]      FIG. 1B  is a general sectional view through a gas turbine engine along the engine longitudinal axis illustrating an engine static structure case arrangement on the lower half thereof; 
           [0028]      FIG. 1C  is a side view of an mount system illustrating a rear mount attached through an engine thrust case to a mid-turbine frame between a first and second bearing supported thereby; 
           [0029]      FIG. 1D  is a forward perspective view of an mount system illustrating a rear mount attached through an engine thrust case to a mid-turbine frame between a first and second bearing supported thereby; 
           [0030]      FIG. 2A  is a top view of an engine mount system; 
           [0031]      FIG. 2B  is a side view of an engine mount system within a nacelle system; 
           [0032]      FIG. 2C  is a forward perspective view of an engine mount system within a nacelle system; 
           [0033]      FIG. 3  is a side view of an engine mount system within another front mount; 
           [0034]      FIG. 4A  is an aft perspective view of an aft mount; 
           [0035]      FIG. 4B  is an aft view of an aft mount of  FIG. 4A ; 
           [0036]      FIG. 4C  is a front view of the aft mount of  FIG. 4A ; 
           [0037]      FIG. 4D  is a side view of the aft mount of  FIG. 4A ; 
           [0038]      FIG. 4E  is a top view of the aft mount of  FIG. 4A ; 
           [0039]      FIG. 5A  is a side view of the aft mount of  FIG. 4A  in a first slide position; 
           [0040]    and 
           [0041]      FIG. 5B  is a side view of the aft mount of  FIG. 4A  in a second slide position. 
           [0042]      FIG. 6  shows another embodiment. 
           [0043]      FIG. 7  shows yet another embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0044]      FIG. 1A  illustrates a general partial fragmentary schematic view of a gas turbofan engine  10  suspended from an engine pylon  12  within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation. 
         [0045]    The turbofan engine  10  includes a core engine within a core nacelle C that houses a low spool  14  and high spool  24 . The low spool  14  includes a low pressure compressor  16  and low pressure turbine  18 . The low spool  14  drives a fan section  20  connected to the low spool  14  either directly or through a gear train  25 . 
         [0046]    The high spool  24  includes a high pressure compressor  26  and high pressure turbine  28 . A combustor  30  is arranged between the high pressure compressor  26  and high pressure turbine  28 . The low and high spools  14 ,  24  rotate about an engine axis of rotation A. 
         [0047]    The engine  10  in one non-limiting embodiment is a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine  10  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the gear train  25  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  18  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  10  bypass ratio is greater than ten (10:1), the turbofan diameter is significantly larger than that of the low pressure compressor  16 , and the low pressure turbine  18  has a pressure ratio that is greater than 5:1. The gear train  25  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 2.5:1. 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. 
         [0048]    Airflow enters the fan nacelle F which at least partially surrounds the core nacelle C. The fan section  20  communicates airflow into the core nacelle C to the low pressure compressor  16 . Core airflow compressed by the low pressure compressor  16  and the high pressure compressor  26  is mixed with the fuel in the combustor  30  where is ignited, and burned. The resultant high pressure combustor products are expanded through the high pressure turbine  28  and low pressure turbine  18 . The turbines  28 ,  18  are rotationally coupled to the compressors  26 ,  16  respectively to drive the compressors  26 ,  16  in response to the expansion of the combustor product. The low pressure turbine  18  also drives the fan section  20  through gear train  25 . A core engine exhaust E exits the core nacelle C through a core nozzle  43  defined between the core nacelle C and a tail cone  33 . 
         [0049]    With reference to  FIG. 1B , the low pressure turbine  18  includes a low number of stages, which, in the illustrated non-limiting embodiment, includes three turbine stages,  18 A,  18 B,  18 C. The gear train  22  operationally effectuates the significantly reduced number of stages within the low pressure turbine  18 . The three turbine stages,  18 A,  18 B,  18 C facilitate a lightweight and operationally efficient engine architecture. It should be appreciated that a low number of stages contemplates, for example, three to six (3-6) stages. Low pressure turbine  18  pressure ratio is pressure measured prior to inlet of low pressure turbine  18  as related to the pressure at the outlet of the low pressure turbine  18  prior to exhaust nozzle. 
         [0050]    Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. The Variable Area Fan Nozzle (“VAFN”)  42  operates to effectively vary the area of the fan nozzle exit area  44  to selectively adjust the pressure ratio of the bypass flow B in response to a controller C. Low pressure ratio turbofans are desirable for their high propulsive efficiency. However, low pressure ratio fans may be inherently susceptible to fan stability/flutter problems at low power and low flight speeds. The VAFN  42  allows the engine to change to a more favorable fan operating line at low power, avoiding the instability region, and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise. 
         [0051]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  20  of the engine  10  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 (‘TSFCT’)”—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 the Fan Exit Guide Vane (“FEGV”) system  36 . 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 [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
         [0052]    As the fan blades within the fan section  20  are efficiently designed at a particular fixed stagger angle for an efficient cruise condition, the VAFN  42  is operated to effectively vary the fan nozzle exit area  44  to adjust fan bypass air flow such that the angle of attack or incidence on the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff to thus provide optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels. 
         [0053]    The engine static structure  44  generally has sub-structures including a case structure often referred to as the engine backbone. The engine static structure  44  generally includes a fan case  46 , an intermediate case (IMC)  48 , a high pressure compressor case  50 , a combustor case  52 A, a high pressure turbine case  52 B, a thrust case  52 C, a low pressure turbine case  54 , and a turbine exhaust case  56  ( FIG. 1B ). Alternatively, the combustor case  52 A, the high pressure turbine case  52 B and the thrust case  52 C may be combined into a single case. It should be understood that this is an exemplary configuration and any number of cases may be utilized. 
         [0054]    The fan section  20  includes a fan rotor  32  with a plurality of circumferentially spaced radially outwardly extending fan blades  34 . The fan blades  34  are surrounded by the fan case  46 . The core engine case structure is secured to the fan case  46  at the IMC  48  which includes a multiple of circumferentially spaced radially extending struts  40  which radially span the core engine case structure and the fan case  20 . 
         [0055]    The engine static structure  44  further supports a bearing system upon which the turbines  28 ,  18 , compressors  26 ,  16  and fan rotor  32  rotate. A # 1  fan dual bearing  60  which rotationally supports the fan rotor  32  is axially located generally within the fan case  46 . The # 1  fan dual bearing  60  is preloaded to react fan thrust forward and aft (in case of surge). A # 2  LPC bearing  62  which rotationally supports the low spool  14  is axially located generally within the intermediate case (IMC)  48 . The # 2  LPC bearing  62  reacts thrust. A # 3  fan dual bearing  64  which rotationally supports the high spool  24  and also reacts thrust. The # 3  fan bearing  64  is also axially located generally within the IMC  48  just forward of the high pressure compressor case  50 . A # 4  bearing  66  which rotationally supports a rear segment of the low spool  14  reacts only radial loads. The # 4  bearing  66  is axially located generally within the thrust case  52 C in an aft section thereof. A # 5  bearing  68  rotationally supports the rear segment of the low spool  14  and reacts only radial loads. The # 5  bearing  68  is axially located generally within the thrust case  52 C just aft of the # 4  bearing  66 . It should be understood that this is an exemplary configuration and any number of bearings may be utilized. 
         [0056]    The # 4  bearing  66  and the # 5  bearing  68  are supported within a mid-turbine frame (MTF)  70  to straddle radially extending structural struts  72  which are preloaded in tension ( FIGS. 1C-1D ). The MTF  70  provides aft structural support within the thrust case  52 C for the # 4  bearing  66  and the # 5  bearing  68  which rotatably support the spools  14 ,  24 . 
         [0057]    A dual rotor engine such as that disclosed in the illustrated embodiment typically includes a forward frame and a rear frame that support the main rotor bearings. The intermediate case (IMC)  48  also includes the radially extending struts  40  which are generally radially aligned with the # 2  LPC bearing  62  ( FIG. 1B ). It should be understood that various engines with various case and frame structures will benefit from the present invention. 
         [0058]    The turbofan gas turbine engine  10  is mounted to aircraft structure such as an aircraft wing through a mount system  80  attachable by the pylon  12 . The mount system  80  includes a forward mount  82  and an aft mount  84  ( FIG. 2A ). The forward mount  82  is secured to the IMC  48  and the aft mount  84  is secured to the MTF  70  at the thrust case  52 C. The forward mount  82  and the aft mount  84  are arranged in a plane containing the axis A of the turbofan gas turbine  10 . This eliminates the thrust links from the intermediate case, which frees up valuable space beneath the core nacelle and minimizes IMC  48  distortion. 
         [0059]    Referring to  FIGS. 2A-2C , the mount system  80  reacts the engine thrust at the aft end of the engine  10 . The term “reacts” as utilized in this disclosure is defined as absorbing a load and dissipating the load to another location of the gas turbine engine  10 . 
         [0060]    The forward mount  82  supports vertical loads and side loads. The forward mount  82  in one non-limiting embodiment includes a shackle arrangement which mounts to the IMC  48  at two points  86 A,  86 B. The forward mount  82  is generally a plate-like member which is oriented transverse to the plane which contains engine axis A. Fasteners are oriented through the forward mount  82  to engage the intermediate case (IMC)  48  generally parallel to the engine axis A. In this illustrated non-limiting embodiment, the forward mount  82  is secured to the IMC  40 . In another non-limiting embodiment, the forward mount  82  is secured to a portion of the core engine, such as the high-pressure compressor case  50  of the gas turbine engine  10  (see  FIG. 3 ). One of ordinary skill in the art having the benefit of this disclosure would be able to select an appropriate mounting location for the forward mount  82 . 
         [0061]    Referring to  FIG. 4A , the aft mount  84  generally includes a first A-arm  88 A, a second A-arm  88 B, a rear mount platform  90 , a whiffle tree assembly  92  and a drag link  94 . The rear mount platform  90  is attached directly to aircraft structure such as the pylon  12 . The first A-arm  88 A and the second A-arm  88 B mount between the thrust case  52 C at case bosses  96  which interact with the MTF  70  ( FIGS. 4B-4C ), the rear mount platform  90  and the whiffle tree assembly  92 . It should be understood that the first A-arm  88 A and the second A-arm  88 B may alternatively mount to other areas of the engine  10  such as the high pressure turbine case or other cases. It should also be understood that other frame arrangements may alternatively be used with any engine case arrangement. 
         [0062]    Referring to  FIG. 4D , the first A-arm  88 A and the second A-arm  88 B are rigid generally triangular arrangements, each having a first link arm  89   a , a second link arm  89   b  and a third link arm  89   c . The first link arm  89   a  is between the case boss  96  and the rear mount platform  90 . The second link arm  89   b  is between the case bosses  96  and the whiffle tree assembly  92 . The third link arm  89   c  is between the whiffle tree assembly  92  rear mount platform  90 . The first A-arm  88 A and the second A-arm  88 B primarily support the vertical weight load of the engine  10  and transmit thrust loads from the engine to the rear mount platform  90 . 
         [0063]    The first A-arm  88 A and the second A-arm  88 B of the aft mount  84  force the resultant thrust vector at the engine casing to be reacted along the engine axis A which minimizes tip clearance losses due to engine loading at the aft mount  84 . This minimizes blade tip clearance requirements and thereby improves engine performance. 
         [0064]    The whiffle tree assembly  92  includes a whiffle link  98  which supports a central ball joint  100 , a first sliding ball joint  102 A and a second sliding ball joint  102 B ( FIG. 4E ). It should be understood that various bushings, vibration isolators and such like may additionally be utilized herewith. The central ball joint  100  is attached directly to aircraft structure such as the pylon  12 . The first sliding ball joint  102 A is attached to the first A-arm  88 A and the second sliding ball joint  102 B is mounted to the first A-arm  88 A. The first and second sliding ball joint  102 A,  102 B permit sliding movement of the first and second A-arm  88 A,  88 B (illustrated by arrow S in  FIGS. 5A and 5B ) to assure that only a vertical load is reacted by the whiffle tree assembly  92 . That is, the whiffle tree assembly  92  allows all engine thrust loads to be equalized transmitted to the engine pylon  12  through the rear mount platform  90  by the sliding movement and equalize the thrust load that results from the dual thrust link configuration. The whiffle link  98  operates as an equalizing link for vertical loads due to the first sliding ball joint  102 A and the second sliding ball joint  102 B. As the whiffle link  98  rotates about the central ball joint  100  thrust forces are equalized in the axial direction. The whiffle tree assembly  92  experiences loading only due to vertical loads, and is thus less susceptible to failure than conventional thrust-loaded designs. 
         [0065]    The drag link  94  includes a ball joint  104 A mounted to the thrust case  52 C and ball joint  104 B mounted to the rear mount platform  90  ( FIGS. 4B-4C ). The drag link  94  operates to react torque. 
         [0066]    The aft mount  84  transmits engine loads directly to the thrust case  52 C and the MTF  70 . Thrust, vertical, side, and torque loads are transmitted directly from the MTF  70  which reduces the number of structural members as compared to current in-practice designs. 
         [0067]    The mount system  80  is compact, and occupies space within the core nacelle volume as compared to turbine exhaust case-mounted configurations, which occupy space outside of the core nacelle which may require additional or relatively larger aerodynamic fairings and increase aerodynamic drag and fuel consumption. The mount system  80  eliminates the heretofore required thrust links from the IMC, which frees up valuable space adjacent the IMC  48  and the high pressure compressor case  50  within the core nacelle C. 
         [0068]    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. 
         [0069]      FIG. 6  shows an embodiment  200 , wherein there is a fan drive turbine  208  driving a shaft  206  to in turn drive a fan rotor  202 . A gear reduction  204  may be positioned between the fan drive turbine  208  and the fan rotor  202 . This gear reduction  204  may be structured and operate like the gear reduction disclosed above. A compressor rotor  210  is driven by an intermediate pressure turbine  212 , and a second stage compressor rotor  214  is driven by a turbine rotor  216 . A combustion section  218  is positioned intermediate the compressor rotor  214  and the turbine section  216 . 
         [0070]      FIG. 7  shows yet another embodiment  300  wherein a fan rotor  302  and a first stage compressor  304  rotate at a common speed. The gear reduction  306  (which may be structured as disclosed above) is intermediate the compressor rotor  304  and a shaft  308  which is driven by a low pressure turbine section. 
         [0071]    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.