Abstract:
A fan assembly for a gas turbine engine includes a fan including a plurality of fan blades extending radially outwardly from a fan hub to a blade tip defining a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath having a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle. The throat portion has a minimum throat radius measured from a fan central axis greater than a blade tip radius.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of an earlier filing date from U.S. Provisional Application Ser. No. 62/273,560 filed Dec. 31, 2015, the entire disclosure of which is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    This disclosure relates to gas turbine engines, and more particularly to nacelle inlets for gas turbine engines. 
         [0003]    Gas turbine engines, such as those used to power modern commercial and military aircraft, generally includes a fan section where an airflow is introduced to the gas turbine engine, a compressor section to pressurize the airflow, a combustor section for burning hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. The airflow flows along a gaspath through the gas turbine engine. The gas turbine engine is typically enclosed in a nacelle, with the airflow introduced to the gas turbine engine at the fan via a nacelle inlet. 
         [0004]    The nacelle inlet is utilized to smooth the airflow into the gas turbine engine, and is designed to reduce the airflow speed from upstream cruise speed of 0.8 Mach down to about 0.5 Mach. The typical nacelle inlet cross-section includes a throat at which the inlet diameter is at its narrowest and a diffuser section downstream of the throat, relative to the general flow direction of the airflow. The diffuser section typically slows the airflow just upstream of the fan. 
         [0005]    A typical nacelle inlet is characterized by an inlet length (L), which is an axial length from a forwardmost point of the nacelle to a leading edge blade tip of the fan and a diameter (D) between opposing leading edge blade tips inlet. A conventional nacelle inlet has a ratio L/D of about 0.5. As gas turbine engine designs have evolved, designs have moved toward larger diameter fans, which require a greater inlet length to attain the L/D of 0.5. This increase in axial length of the nacelle increases overhang, adds weight and increases drag due to the nacelle, with all of these negatively affecting performance of the aircraft on which the gas turbine engine is utilized. To alleviate these issues, shorter inlet lengths have been proposed, increasing the amount of diffusion required per unit length to slow the incoming airflow to the desired 0.5 Mach. 
         [0006]    It is often desired or necessary to remove fan blades from the fan for repair or replacement. It is desired to do this removal through the inlet with the inlet installed to minimize disassembly of the gas turbine engine. The shorter inlets, however, result in an inlet throat that is at a smaller radius than the fan radius, and because of the small throat radius and the nearness of the throat to the fan blades, removal of the fan blades may not be possible without removal of the inlet from the gas turbine engine. 
       SUMMARY 
       [0007]    In one embodiment, a fan assembly for a gas turbine engine includes a fan including a plurality of fan blades. Each fan blade extends radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath. The nacelle flowpath has a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle. The throat portion has a minimum throat radius measured from a fan central axis greater than a blade tip radius measured relative to the fan central axis. 
         [0008]    Additionally or alternatively, in this or other embodiments a ratio of a fan hub outer radius to an inner surface radius measured from the central axis at an axial position corresponding to a forwardmost fan blade tip at the topmost portion of the nacelle is greater than the ratio of fan hub outer radius to inner surface radius at a remainder of a nacelle perimeter. 
         [0009]    Additionally or alternatively, in this or other embodiments the ratio of fan hub outer radius to inner surface radius at the topmost portion is in the range of 0.35 to 0.40. 
         [0010]    Additionally or alternatively, in this or other embodiments the topmost portion is defined as a sixty circumferential degree portion of the nacelle centered on a top dead center of the nacelle. 
         [0011]    Additionally or alternatively, in this or other embodiments the fan assembly has a fan pressure ratio between 1.2 and 1.45. 
         [0012]    Additionally or alternatively, in this or other embodiments the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter. A ratio of the hub diameter to the fan diameter is in the range of 0.25 to 0.45. 
         [0013]    Additionally or alternatively, in this or other embodiments a spinner is positioned at a central fan axis further defining the nacelle flowpath between the spinner and the nacelle inner surface. 
         [0014]    Additionally or alternatively, in this or other embodiments a minimum distance between the spinner and the nacelle inner surface is greater than a fan blade radial length. 
         [0015]    Additionally or alternatively, in this or other embodiments one or more fan blades are removable via the fan inlet without removal of the nacelle. 
         [0016]    In another embodiment a gas turbine engine includes a turbine and a fan assembly operably connected to the turbine. The fan assembly includes a fan including a plurality of fan blades. Each fan blade extends radially outwardly from a fan hub to a blade tip, the plurality of blade tips defining a fan diameter. A nacelle surrounds the fan and defining a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades, a ratio of inlet length to fan diameter between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath. The nacelle flowpath has a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle, the throat portion having a minimum throat radius measured from a fan central axis greater than a blade tip radius measured relative to the fan central axis. 
         [0017]    Additionally or alternatively, in this or other embodiments a ratio of a fan hub outer radius to an inner surface radius measured from the central axis at an axial position corresponding to a forwardmost fan blade tip at the topmost portion of the nacelle is greater than the ratio of fan hub outer radius to inner surface radius at a remainder of a nacelle perimeter. 
         [0018]    Additionally or alternatively, in this or other embodiments the ratio of fan hub outer radius to inner surface radius at the topmost portion is in the range of 0.35 to 0.40. 
         [0019]    Additionally or alternatively, in this or other embodiments the topmost portion is defined as a sixty circumferential degree portion of the nacelle centered on a top dead center of the nacelle. 
         [0020]    Additionally or alternatively, in this or other embodiments the fan assembly has a fan pressure ratio between 1.2 and 1.45. 
         [0021]    Additionally or alternatively, in this or other embodiments the plurality of fan blades extend radially outwardly from a fan hub at a hub diameter, a ratio of the hub diameter to the fan diameter in the range of 0.25 to 0.45. 
         [0022]    Additionally or alternatively, in this or other embodiments a spinner is positioned at a central fan axis further defining the nacelle flowpath between the spinner and the nacelle inner surface. 
         [0023]    Additionally or alternatively, in this or other embodiments a minimum distance between the spinner and the nacelle inner surface is greater than a fan blade radial length. 
         [0024]    Additionally or alternatively, in this or other embodiments one or more fan blades are removable via the fan inlet without removal of the nacelle. 
         [0025]    In yet another embodiment, a nacelle for a fan of a gas turbine engine includes a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of a fan, a ratio of inlet length to a fan diameter between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath. The nacelle flowpath has a convex throat portion and a diffusion portion between the throat portion and the leading edge of the fan blade at a topmost portion of the nacelle. The throat portion has a minimum throat radius measured from a fan central axis greater than a blade tip radius of a fan blade measured relative to the fan central axis. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]    The subject matter which is regarded as the present disclosure is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
           [0027]      FIG. 1  illustrates a schematic cross-sectional view of an embodiment of a gas turbine engine; 
           [0028]      FIG. 2  illustrates a cross-sectional view of an embodiment of a fan of a gas turbine engine; and 
           [0029]      FIG. 3  illustrates a cross-section view of a top half of an embodiment of a fan for a gas turbine engine. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines may include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air along a core flow path C where air is compressed and communicated to the combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0031]    Although the disclosed non-limiting embodiment depicts a two-spool turbofan gas turbine engine, 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, for example, a turbine engine including three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0032]    The example gas turbine 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. 
         [0033]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such 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 (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
         [0034]    The combustor section  26  includes a combustor  56  arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0035]    The example low pressure turbine  46  has a pressure ratio that is greater than about 5. The pressure ration of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0036]    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  as well as settling airflow entering the low pressure turbine  46 . 
         [0037]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited at the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vanes  59  of the mid-turbine frame  57  as the inlet guide vane for the low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  57 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  28  is increased and a higher power density may be achieved. 
         [0038]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0039]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0040]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and at an altitude of about 35,000 feet. The flight condition of 0.8 Mach and 35,000 feet, with the gas turbine engine  20  at its optimal fuel efficiency—also known as “bucket cruise Thrust Specific Fuel Consumption (TSFC)”—is an industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the gas turbine engine  20  produces at that minimum point. 
         [0041]    “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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. In some embodiments, the low fan pressure ratio is between 1.2 and 1.45. 
         [0042]    “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 nonlimiting embodiment, is less than about 1150 ft/sec. 
         [0043]      FIG. 2  illustrates an example embodiment of the gas turbine engine  20  with a nacelle  80  or cowling, that surrounds the entire gas turbine engine  20 . An inlet portion  82  is situated forward of the fan  42 . In this example, the inlet portion  82  has a leading edge  84 , which may be defined by an inlet side cut in the nacelle  80 . The leading edge  84  is generally within a first reference plane  86 . 
         [0044]    The nacelle  80  in some examples includes a flange  87  that is received against a leading edge of a fan case  88 . The inlet portion has a length L between a selected location corresponding to the leading edge  84 , such as a location within the first reference plane  86 , and a forwardmost portion  90  of leading edges of fan blades  92  of the fan  42 . In this example, the length L may be considered an axial length of the inlet portion  82  because the length L is measured along a direction parallel to the central longitudinal axis A of the gas turbine engine  20 . In the illustrated example, the inlet portion  82  of the nacelle  80  and the section of the fan case  88  that is forward of the fan blades  92  collectively establish the length. In other words, in this example, the length L of the inlet portion  82  includes the length of the inlet section of the nacelle  80  and a portion of the length of the fan case  88 . 
         [0045]    The fan blades  92  may be unswept as shown in  FIG. 1 , or may alternatively be swept fan blades  92  as shown in  FIG. 2 . In some examples, the fan blades  92  are conventional radial fan blades  92  or three-dimensionally swept fan blades  92 . In other embodiments, the fan blades  92  may be forward-swept fan blades  92  or rearward-swept fan blades  92 . In still other embodiments, the fan blades  92  may include a combination of forward sweep and rearward sweep. 
         [0046]    The fan blades  92  establish a diameter between circumferentially outermost edges, or blade tips  94 . The fan diameter D is shown in  FIG. 2  as a dimension extending between the blade tips  94  of two of the fan blades  92  that are parallel to each other and extending in opposite directions from the central axis A. In the illustration, the forwardmost portions  90  of the fan blades  92  are within a second reference plane  96 . In this example, the second reference plane  96  is oriented generally perpendicular to the central axis A. The first reference plane  86  in this example is oriented at an oblique angle relative to the second reference plane  96  and the central axis A. In the illustrated example, the oblique angle of orientation of the first reference plane  86  relative to the second reference plane  96  is about 5 degrees. 
         [0047]    The length L is selected to establish a desired dimensional relationship between L and D. In some embodiments, the dimension relationship of L/D (e.g. the ratio of L/D) is between about 0.20 and 0.45. In some examples, L/D is between about 0.30 and about 0.40. In some embodiments, the ratio L/D is about 0.35. 
         [0048]    As can be appreciated from  FIG. 2 , the length L of the inlet portion  82  is different at different locations along the perimeter of the nacelle  80 . The second reference plane  96  is further from the first reference plane  86  nearest the top (according to the drawing) of the gas turbine engine  20  than it is nearest the bottom of the gas turbine engine  20 . Thus, L is greater nearest the top of the gas turbine engine  20  than it is nearest the bottom of the gas turbine engine  20 . The greatest length L in this example corresponds to a value of L/D that is no greater than about 0.45. The smallest length L in the illustrated example corresponds to a value of L/D that is at least about 0.20. The value of L/D may vary between those two values at different circumferential locations around the gas turbine engine  20 . In one example where first reference plane  86  has a variable distance from the second reference plane  96 , the dimensional relationship L/D is based on an average distance between the first reference plane  86  and the second reference plane  96 . 
         [0049]    Shown in  FIG. 3  is another cross-sectional view of the nacelle  80  and fan  42  of the gas turbine engine  20 . The nacelle  80  includes an inner surface  100  that, together with a spinner  102  located at central axis A, defines inlet flowpath  104  for bypass airflow B and core airflow C. More particularly, inner surface  100  of the nacelle  80  defines an outer flowpath surface, while spinner  102  defines an inner flowpath surface, relative to the central axis A. The fan blades  92  extend outwardly from a blade root  106  at fan hub  95  to the blade tip  94 . The fan hub  95  defines hub radius  110  at the blade root  106 , as measured from central axis A, while the blade tips  94  define a tip radius  112 . In some embodiments, a ratio of the hub radius  110  to the tip radius  112  is in the range of 0.25 to 0.45. 
         [0050]    The nacelle  80  inner surface  100  at an upper portion  120  of the nacelle  80 , for example a 60 degree circumferential portion centered on a top dead center  118  of the nacelle  80 , has a throat portion  122  with a minimum throat radius  124  measured from the central axis A, that is greater than the tip radius  112 . In some embodiments, a minimum distance between spinner  102  and the inner surface  100  of nacelle  80  is greater than a fan blade length  126 , or distance between the blade root  106  and blade tip  94 . In some embodiments, a ratio of hub radius  110  to an inner surface radius  126  between the fan blade  92  and the throat portion  122  at the upper portion  120  is greater than the ratio of hub radius  110  to inner surface radius  126  at the remainder of the nacelle  80  perimeter. In some embodiments, the ratio of hub radius  110  to inner surface radius  126  is in the range of 0.35 to 0.40 at upper portion  120 , while the ratio of hub radius  110  to inner surface radius is lower than 0.35. Tailoring of the inner surface  100  at upper portion  120  to accommodate removal of the fan blade  92  through the inlet greatly simplifies maintenance of the gas turbine engine  10 , while doing such tailoring only at upper portion  120  allows nacelle  80  to still meet performance requirements overall. 
         [0051]    While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Rather, the present disclosure can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the scope of the present disclosure. Additionally, while various embodiments of the present disclosure have been described, it is to be understood that aspects of the present disclosure may include only some of the described embodiments. Accordingly, the present disclosure is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.