Patent Publication Number: US-2022220854-A1

Title: Turbine engine with an airfoil having a set of dimples

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application takes priority to Italian Patent Application Serial No. 102021000000296, filed Jan. 8, 2021. 
     TECHNICAL FIELD 
     This disclosure generally relates to an airfoil for a gas turbine engine, and more specifically to a set of dimples of the airfoil. 
     BACKGROUND 
     Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of compressor and turbine stages, with each stage typically including a complementary set of rotating blades and stationary vanes. Turbine engines include several components that utilize airfoils. By way a of non-limiting example, the airfoils can be located in the engine turbines, compressors, or fans. Airfoils can be defined by an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction. 
     During operation, an airflow can flow over the outer wall of the airfoil in the stream-wise direction from the leading edge toward the trailing edge to define a streamline. On the suction side, the streamline will follow the contour of the airfoil until it separates from the outer wall of the airfoil defining a boundary layer growth. The sooner the boundary layer growth occurs on the suction side of the airfoil, the lower the overall efficiency of the gas turbine engine. 
     BRIEF DESCRIPTION 
     In one aspect the disclosure relates to a gas turbine engine comprising a set of circumferentially adjacent airfoils, the airfoils having an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction, and a set of dimples provided on the outer wall of at least one of the airfoils, with the dimples spaced in at least one of the stream-wise and or span-wise direction, and the dimples having a non-perforated section in a shape defining a two-dimensional surface area that when viewed in planform, the two-dimensional surface area decreases in the stream-wise direction. 
     In another aspect the disclosure relates to an airfoil comprising an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction, and a set of dimples provided on the outer wall, with the dimples spaced in at least one of the stream-wise or span-wise direction, and the dimples having a non-perforated section in a shape defining a two-dimensional surface area that when viewed in planform, the two-dimensional surface area decreases in the stream-wise direction. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present description, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which refers to the appended FIGS., in which: 
         FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine for an aircraft. 
         FIG. 2  is a is a perspective view of an airfoil assembly of the gas turbine engine of  FIG. 1  having an airfoil defined by a pressure side and a suction side, with a set of dimples provided along a portion of the suction side. 
         FIG. 3  is a schematic top-down view of the airfoil assembly of  FIG. 2 , illustrating a pair of circumferentially adjacent airfoils. 
         FIG. 4  is a perspective view of a dimple of the set of dimples of  FIG. 2  having a trapezoidal two-dimensional surface area when viewed in planform. 
         FIG. 5  is a planform view of the dimple of  FIG. 3 , illustrating a two-dimensional surface area of the dimple. 
         FIG. 6A  is a sectional view of the airfoil of  FIG. 2  taken from cut IV-IV, further illustrating a streamline and a boundary layer growth. 
         FIG. 6B  is a sectional view of an exemplary airfoil taken from cut IV-IV of  FIG. 2 , the exemplary airfoil not include the set of dimples and further illustrating the streamline and the boundary layer growth. 
         FIG. 7  is a planform view of an exemplary dimple of the set of dimples of  FIG. 2 , the exemplary dimple including a triangular two-dimensional surface area. 
         FIG. 8  is a planform view of an exemplary dimple of the set of dimples of  FIG. 2 , the exemplary dimple including a triangular two-dimensional surface area with rounded vertices. 
         FIG. 9  is a planform view of an exemplary dimple of the set of dimples of  FIG. 2 , the exemplary dimple including an exemplary trapezoidal two-dimensional surface area. 
         FIG. 10  is a planform view of an exemplary dimple of the set of dimples of  FIG. 2 , the exemplary dimple including a semi-oval two-dimensional surface area. 
     
    
    
     DETAILED DESCRIPTION 
     Aspects of the disclosure described herein are directed a set of airfoils for a gas turbine engine. The airfoils include an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction. The airfoils can further include a set of dimples provided along a portion of the outer wall on the suction side of a corresponding airfoil. The dimples can extend from a leading edge to a trailing edge in the stream-wise direction and be spaced in at least one of the stream-wise or span-wise direction. Specifically, the dimples can be arranged in rows in the stream-wise direction or columns in the span-wise direction along the outer wall of the airfoil. It will be appreciated that the dimples can be defined by a two-dimensional surface area when viewed in planform that can be formed according to various geometries such as, but not limited to, a triangular two-dimensional surface area, a trapezoidal two-dimensional surface area, a hexagonal two-dimensional surface area, or any other suitable two-dimensional surface area. The dimples can include a first non-perforated section that increases in depth in the stream-wise direction from the leading edge of the dimples to a maximum depth, and a second non-perforated section that decreases in depth from the maximum depth to the trailing edge of the dimples. 
     To reduce the boundary layer growth during operation of the gas turbine engine, the dimples are shaped to generate vortices within the suction side boundary layer. These features counteract the formation of a laminar separation after a throat of the airfoil, thus preventing the separation and boundary layer growth by keeping the flow attached to the surfaced for a greater amount of the airfoil when compared to conventional airfoils without the set of dimples. In other words, the dimples, which generate the vortices, can be used to reduce the effects of the boundary layer growth by avoiding laminar separations on the suction side of the airfoil before reaching the trailing edge. This, in turn, ultimately reduces the total boundary layer growth or a boundary layer thickness, which can result in an increased aerodynamic efficiency of the airfoils within the gas turbine engine. This can ultimately result in an increased power output and efficiency of the gas turbine engine when compared to conventional gas turbine engines that do not include a set of airfoils including the set of dimples. 
     For the purposes of illustration, one exemplary component will be described in the form of a gas turbine engine having a set of airfoils including an outer wall and a set of dimples received on at least a portion of the suctions side of the outer wall. Such a gas turbine engine can be in the form of a gas turbine engine, a turboprop, a turboshaft, a turbofan engine, or a counter rotating turbine (e.g., a gas turbine including an inner and outer rotor in at least a section of the gas turbine engine) in non-limiting examples. It will be understood, however, that aspects of the disclosure described herein are not so limited and can have general applicability with other turbine engine components or within other engine systems. For example, the disclosure can have applicability for systems in other engines or vehicles, and may be used to provide benefits in industrial, commercial, and residential applications. 
     As used herein, the term “upstream” refers to a direction that is opposite the fluid flow direction, and the term “downstream” refers to a direction that is in the same direction as the fluid flow. The term “fore” or “forward” means in front of something and “aft” or “rearward” means behind something. For example, when used in terms of fluid flow, fore/forward can mean upstream and aft/rearward can mean downstream. 
     Additionally, as used herein, the terms “radial” or “radially” refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a center longitudinal axis of the engine and an outer engine circumference. Furthermore, as used herein, the term “set” or a “set” of elements can be any number of elements, including only one. 
     All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader&#39;s understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of aspects of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. 
       FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine  10  for an aircraft. The gas turbine engine  10  has a generally longitudinally extending axis or engine centerline  12  extending forward  14  to aft  16 . The gas turbine engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
     The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a set of fan blades  42  disposed radially about the engine centerline  12 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form an engine core  44  of the gas turbine engine  10 , which generates combustion gases. The engine core  44  is surrounded by core casing  46 , which can be coupled with the fan casing  40 . 
     A HP shaft or spool  48  disposed coaxially about the engine centerline  12  of the gas turbine engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the engine centerline  12  of the gas turbine engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . The spools  48 ,  50  are rotatable about the engine centerline  12  and couple to a set of rotatable elements, which can collectively define a rotor  51 . 
     The LP compressor  24  and the HP compressor  26  respectively include a set of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  (also called a nozzle) to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  can be provided in a ring and can extend radially outwardly relative to the engine centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned upstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  56 ,  58  for a stage of the compressor can be mounted to a disk  61 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  61 . The vanes  60 ,  62  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     The HP turbine  34  and the LP turbine  36  respectively include a set of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine blades  68 ,  70  can be provided in a ring and can extend radially outwardly relative to the engine centerline  12 , from a blade platform to a blade tip, while the corresponding static turbine vanes  72 ,  74  are positioned upstream of and adjacent to the rotating blades  68 ,  70 . It is noted that the number of blades, vanes, and turbine stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  68 ,  70  for a stage of the turbine can be mounted to a disk  71 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having a dedicated disk  71 . The vanes  72 ,  74  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     Complementary to the rotor portion, the stationary portions of the gas turbine engine  10 , such as the static vanes  60 ,  62 ,  72 ,  74  among the compressor and turbine sections  22 ,  32  are also referred to individually or collectively as a stator  63 . As such, the stator  63  can refer to the combination of non-rotating elements throughout the gas turbine engine  10 . 
     In operation, the airflow exiting the fan section  18  is split such that a portion of the airflow is channeled into the LP compressor  24 , which then supplies pressurized airflow  76  to the HP compressor  26 , which further pressurizes the air. The pressurized airflow  76  from the HP compressor  26  is mixed with fuel in the combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and the exhaust gas is ultimately discharged from the gas turbine engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
     A portion of the pressurized airflow  76  can be drawn from the compressor section  22  as bleed air  77 . The bleed air  77  can be drawn from the pressurized airflow  76  and provided to engine components requiring cooling. The temperature of pressurized airflow  76  entering the combustor  30  is significantly increased. As such, cooling provided by the bleed air  77  is necessary for operating of such engine components in the heightened temperature environments. 
     A remaining portion of the airflow  78  bypasses the LP compressor  24  and engine core  44  and exits the gas turbine engine  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a set of airfoil guide vanes  82 , at the fan exhaust side  84 . More specifically, a circumferential row of radially extending airfoil guide vanes  82  are utilized adjacent the fan section  18  to exert some directional control of the airflow  78 . 
     Some of the air supplied by the fan  20  can bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the gas turbine engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . 
     As a non-limiting example,  FIG. 2  is a perspective view of an airfoil assembly  98  received within a portion of the gas turbine engine  10  of  FIG. 1 . The airfoil assembly  98  can include an airfoil  100  extending from a platform  102 , with a dovetail  104  depending from the platform  102 . It will be appreciated that the airfoil assembly  98  can be provided in any suitable portion of the gas turbine engine  10  such as, but not limited to, the compressor section  22  or the turbine section  32 . 
     The airfoil  100  extends from a root  108  to a tip  106  to define a span-wise direction. The airfoil  100  includes an outer wall  110  defining the periphery of the airfoil  100 , which can be functionally divided into a pressure side  112  and a suction side  114 , demarcated by a leading edge  116  and a trailing edge  118 . The airfoil  100  can extend between the leading edge  116  and the trailing edge  118  to define a stream-wise direction. Specifically, a line joining the leading edge  116  to the trailing edge  118 , equidistant from the pressure side  112  and suction side  114  of the outer wall  110  of the airfoils, can define the stream-wise direction. The stream-wise direction may or may not be parallel to the rotational or engine centerline  12  of the gas turbine engine  10 . 
     The airfoil  100  mounts to the platform  102  at the root  108 . The platform can be an annual band extending around the entirety of the engine centerline  12 . A first platform surface  120  can define a surface of the platform  102  in which the airfoil  100  is mounted to at the root  108 . A second platform surface  121  can define a surface of the platform  102  that opposite the first platform surface  120 . It will be appreciated that depending on where the airfoil assembly  98  is mounted, the first and second platform surfaces  120 ,  121  can be further defined by their radial distance to the engine centerline  12 . For example, if the airfoil assembly  98  is provided on the rotor  51  such that the airfoil  100  is defined as one of the turbine blades  68 , then the second platform surface can be defined to be radially closer to the engine centerline  12  than the first platform surface  120 . Conversely, if the airfoil assembly  98  is provided within the gas turbine engine  10  such that the airfoil  100  is defined as at least one of the static turbine vanes  72 , then the first platform surface  120  can be defined to be radially closer to the engine centerline  12  than the second platform surface. 
     A set of cooling holes  122  can be formed in any wall of the component including the outer wall  110  or the platform wall  202  as illustrated. The set of cooling holes  122  can be referencing a single cooling hole or multiple cooling holes. The set of cooling holes  122  can be located by way of non-limiting example, proximate the leading edge  116 , the trailing edge  118  and be located in the platform  102  on the suction side or the pressure side  112  of the airfoil  100 . It should be understood that the locations of the set of cooling holes  122  is for illustrative purposes only and not mean to be limiting. 
     The dovetail  104  can extend from the platform  102  opposite of the airfoil  100 , and can be configured to mount to the any one or more of the platforms, discs, or rotors in the compressor section  22  or turbine section  32  of the gas turbine engine  10 . The dovetail  104  can include a set of inlet passages  124 , illustrated as three inlet passages  124 . It is contemplated that the inlet passages  124  are fluidly coupled to the set of cooling holes  122  to provide a cooling fluid flow (C) for cooling the platform  102 . In another non-limiting example, the inlet passages  124  can provide the cooling fluid flow (C) to an interior of the airfoil  100  for cooling of the airfoil  100 . It should be appreciated that the dovetail  104  is shown in cross-section, such that the inlet passages  124  are housed within the body of the dovetail  104 . 
     At least a portion of the outer wall  110  of the suction side  114  can include a set of dimples  126  arranged in rows extending in the stream-wise direction and columns extending in the span-wise direction. As illustrated, the rows and columns can be parallel to each other, however, it will be appreciated that at least one row or column can be staggered with respect to the others. As such, one or more rows of the set of dimples  126  can begin further upstream or downstream than the rest of the rows of the set of dimples  126 . Similarly, one or more columns of the set of dimples  126  can begin farther radially outward or inward from the root  108  than the rest of the columns. It will be yet further appreciated that each row and column can include the same total number of dimples  126 . It is contemplated, however, that at least one of the rows or columns can include a differing number of dimples  126  than the other rows or columns. 
     The airfoil  100  including the set of dimples  126  as described herein can be provided within any suitable portion of the gas turbine engine  10  such as, but not limited to, the fan section  18 , the compressor section  22 , the combustion section  28 , the turbine section  32  or within an bypass duct (e.g., a duct including at least a portion of the air supplied by the fan  20  that can bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the gas turbine engine  10 ), or any other suitable component. Specifically, the airfoil  100  including the set of dimples  126  can be provided within the LP turbine  36 . It will be further appreciated that while the description herein is related to the airfoil  100 , it can have equal applicability in other engine components that include the set of dimples  126 . Some non-limiting examples of the engine component having including the set of dimples  126  can include the set of compressor blades  56 ,  58 , the set of turbine blades  68 ,  70 , the set of compressor vanes  60 ,  62 , the set of turbine vanes  72 ,  74 , the set of fan blades  42 , the set of outlet guide vanes  82 , a set of inlet guide vanes, a combustor liner, a combustor deflector, a shroud assembly, a guide vane within the air bypass duct, or any other suitable component of the gas turbine engine  10 . 
       FIG. 3  is a schematic top-down view of the airfoil assembly  98  of  FIG. 2  including a set of circumferentially adjacent airfoils  100 , specifically a pair of circumferentially-adjacent airfoils  100  extending from the first platform surface  120  of the platform  102 . Although the airfoil assembly  98  is illustrated as including two airfoils  100  that are circumferentially-adjacent one another, it will be appreciated that there can be any number of one or more circumferentially adjacent airfoils  100 . 
     A throat  152  can be defined by the minimum distance between the pair of circumferentially-adjacent airfoils  100  within the gas turbine engine  10 . Specifically, the throat  152  can be between the outer wall  110  of one of the airfoils  100  and the trailing edge  118  of the other airfoil  100 , as illustrated. 
     As illustrated, both of the airfoils  100  can include the set of dimples  126  provided along a portion of the outer wall  110  on the suction side  114  of each respective airfoil  100 . It will be appreciated that each airfoil  100  of the airfoil assembly  98  can include the set of dimples  126  provided a length along the outer wall  110  on the suction side  114  of each corresponding airfoil  100 . Alternatively, it is contemplated that any number of one or more airfoils  100  of the airfoil assembly  98  can include the set of dimples  126 . 
       FIG. 4  illustrates a dimple  126  taken from view IV-IV of  FIG. 2 . It will be appreciated that although described in terms of the dimple  126 , that the description of the dimple  126  can be applied to any dimple in the set of dimples  126  provided along the suction side  114  of the airfoil  100 . 
     As illustrated, the dimple  126  can be included on a portion of the airfoil  100  exposed to a working airflow, specifically a streamline  150  that extends in the stream-wise direction. The dimple  126  can include a leading edge  128  and a trailing edge  130 , downstream of the leading edge  128  in the stream-wise direction, with the streamline  150  extending between the leading edge  128  and the trailing edge  130 . The leading edge  128  and the trailing edge  130  of the dimple can be joined by a first lateral edge  132  and a second lateral edge  134  provided on opposing span-wise portions of the dimple  126 . As illustrated, but not necessary, the first and second lateral edges  132 ,  134  can be mirrored-images of one another in the streamwise direction such that they can extend between the leading edge  128  and the trailing edge  130  at the same rate. 
     The dimple  126  can extend below a surface of the airfoil  100  defined by the contour of the outer wall  110  of the airfoil  100 . The portion of the dimple  126  that extends below the surface can define a non-perforated section  136  that extends from the leading edge  128  to the trailing edge  130  and between the first and second lateral edges  132 ,  134 . As used herein, the term “non-perforated” can refer to a section of the airfoil  100 , specifically the dimples  126 , that does not include a through hole (e.g., cooling hole  122 , film hole, etc.) that leads into an internal portion of the airfoil  100  or platform  102 . In other words, the non-perforated section  136  does not include perforations and is formed as a solid, continuous section of the outer wall  110 . The non-perforated section  136  can define a first bottom wall  138 , a second bottom wall  140 , a first lateral wall  142 , and a second lateral wall  144 . The first and second lateral walls  142 ,  144  can each extend from the first or second bottom walls  138 ,  140  to the first and second lateral edges  132 ,  134 , respectively. The first and second bottom walls  138 ,  140 , however, can extend between the leading edge  128  to the trailing edge  130  in the stream-wise direction. It is contemplated that the first and second bottom walls  138 ,  140  can each vary in depth in the stream-wise direction. Specifically, the first bottom wall  138  can extend from the leading edge  128  to a maximum depth  146 , and the second bottom wall  140  can extend from the maximum depth  146  to the trailing edge  130 . As such, the non-perforated section  136  can define a depth of the dimple  126 , which first increases from the leading edge  128  to the maximum depth  146  and then decrease from the maximum depth  146  to the trailing edge  130 . It will be appreciated that the maximum depth  146  can extend across the entire span-wise length of the dimple  126 . Alternatively, the maximum depth  146  can be defined by a single point within a portion of the non-perforated section  136 . 
     It is contemplated that the first bottom wall  138  can continuously increase in depth from the leading edge  128  to the maximum depth  146 . As such, the first bottom wall  138  of the non-perforated section  136  can be defined as a continuously decreasing, or linearly decreasing section of the non-perforated section  136 . Alternatively, at least a portion of the first bottom wall  138  can be non-continuous, or non-linear. For example, the first bottom wall  138  can define a curved cross section when viewed in the span-wise direction. Alternatively, the first bottom wall  138  can be formed by as a step function such that the second bottom wall  140  can include a sudden or instantaneous decrease in depth in the span-wise direction. 
     It is contemplated that the second bottom wall  140  can define a curved cross section when viewed in the span-wise direction. In other words, the second bottom wall  140  can non-continuously, or non-linearly extend from the maximum depth  146  to the trailing edge  130 . As such, the second bottom wall  140  of the non-perforated section  136  can be defined as a non-continuously increasing, or non-linearly increasing section of the non-perforated section  136 . As illustrated, the second bottom wall  140  can define an exponential curve increasing from the maximum depth  146  to the trailing edge  130 . Alternatively, at least a portion of the second bottom wall  140  can be continuous or linearly decreasing in depth from the maximum depth  146  to the trailing edge  130 . 
     The first and second lateral walls  142 ,  144  of the non-perforated section  136  can each be normal to the span-wise direction. Alternatively, at least a portion of the first and second lateral walls  142 ,  144  can decrease in depth, continuously, linearly, non-continuously or non-linearly, from a corresponding portion of the first or second bottom wall  140 ,  140  to the first or second lateral edges  132 ,  134 , respectively. 
       FIG. 5  is a planform view of the dimple  126  of  FIG. 4  illustrating a two-dimensional surface area of the dimple  126  demarcated by the leading edge  128 , the trailing edge  130 , the first lateral edge  132 , and the second lateral edge  134 . 
     It is contemplated that the first and second bottom walls  138 ,  140  can each be defined by non-equal lengths in the stream-wise direction. Specifically, the length of the first bottom wall  138  can be larger than the length of the second bottom wall  140 . As such, the maximum depth  146  can be provided at a portion of the dimple  126  that is greater than 50% of the total length in the stream-wise direction. Specifically, the maximum depth  146  can be provided between 50% or greater than 50% and less than 100% of the stream-wise length of the dimple  126 . It is contemplated, however, that the maximum depth  146  can be provided along a portion of the stream-wise length of the dimple that is greater than 50% of the stream-wise length. In other words, the maximum depth  146  can be provided downstream a midpoint of the dimple  126  in the stream-wise direction. Alternatively, the maximum depth  146  can be provided along any portion of the stream-wise length of the dimple  126 . It will be yet further appreciated that the depth of the maximum depth  146  can be between 2% and 50% of the stream-wise length of the dimple  126 . 
     The non-perforated section  136  of the dimple  126  can define the two-dimensional surface area, specifically a trapezoidal two-dimensional surface area. As illustrated, the first and second lateral edges  132 ,  134  can be angled such that they converge from the leading edge  128  toward the trailing edge  130 . In other words, the two-dimensional surface area can decrease in the stream-wise direction. Specifically, the two-dimensional surface area can decrease from the leading edge  128  to the trailing edge  130 . 
       FIGS. 6A-6B  illustrates a non-limiting example of the airfoil  100  and an exemplary airfoil  200 , with the airfoil  200  being similar to the airfoil  100 , however, the airfoil  200  does not include the set of dimples  126 . The airfoil  200  is similar to the airfoil  100 ; therefore, like parts will be identified with like numerals in the  200  series, with it being understood that the description of the like parts of the airfoil  100  applies to the airfoil  200  unless otherwise noted.  FIGS. 6A-6B  illustrate a comparison between the two airfoils  100 ,  200 , specifically  FIGS. 6A-6B  illustrate the streamline  150  flowing over a portion of the airfoil  100 ,  200  until it separates from the airfoil  100 ,  200  to define a boundary layer growth  154 ,  254 , respectively, with the boundary layer growth  254  of the airfoil  200  beginning sooner along the airfoil  200 , and thus being larger than the boundary layer growth  154  of the airfoil  100 . 
     By way of non-limiting example,  FIG. 6A  illustrates a sectional view of the airfoil  100  taken along sight line IV-IV of  FIG. 2 , further illustrating an airflow  148  defining a suction side  114  streamline  150  that flows over the outer wall  110  of the airfoil  100  from the leading edge  116  toward the trailing edge  118 . The airflow  148  can be any suitable airflow  148  within the gas turbine engine  10 . For example, the airflow  148  can be a working airflow within the gas turbine engine  10  configured to be drawn in through (e.g., within the fan section  18 , compressed by (e.g., within the compressor section  22 ), combusted within (e.g. within the combustion section  28 ), rotate (e.g., within the turbine section  32 ), or otherwise be exhausted through (e.g., within the exhaust section  38 ) at least a section of the gas turbine engine  10 . Alternatively, the airflow  148  can be a bypass airflow within the bypass duct of the gas turbine engine  10 . 
     As discussed herein, the airfoil  100  can be included within a set of circumferentially adjacent airfoils  100  and the throat  152  can be defined between circumferentially-adjacent airfoils  100 . As illustrated, at least one row of the set of dimples  126  can begin upstream the throat  152 . Additionally, or alternatively, at least one row of the set of dimples  126  can be provided along a portion of the outer wall  110  at or downstream the throat  152 . 
     As illustrated, the set of dimples  126  can end proximate the trailing edge  118  of the airfoil  100 . As such, the total amount of the airfoil  100  that includes the set of dimples  126  can be between 30% and 98% of a stream-wise length of the airfoil  100  along the outer wall  110  on the suction side  114 . 
       FIG. 6B  illustrates a sectional view of the airfoil  200  taken from cut IV-IV of  FIG. 2 . The airfoil  200  does not include the set of dimples  126  as the airfoil  100  does. The airflow  148  can define a suction side  114  streamline  250  that flows over the outer wall  210  of the airfoil  200  from the leading edge  116  toward the trailing edge  118 . 
     In operation of the gas turbine engine  10 , at least a portion of the airflow  148  can impinge and be separate by the leading edge  116  of the airfoil  100 ,  200 . The portion of the airflow  148  that flows over the suction side  114  can define the streamline  150 ,  250 , which follows a contour defined by the outer wall  110 ,  210  until it separates from the outer wall  110 ,  210  at a downstream section of the airfoil  100 ,  200 . This separation can define a boundary layer growth  154 ,  254 . 
       FIG. 6A  further illustrates an enlarged view  151  of a portion of the airfoil  100 . Specifically, illustrates the dimple  126  that extends in the stream-wise direction from the leading edge  128  to the trailing edge  130 , and includes the non-perforated section  136  with the first bottom wall  138 , which extends below a surface of the airfoil  100  from the leading edge  128  to the maximum depth  146  and a second bottom wall  140  that extends back to the surface of the airfoil  100  from the maximum depth  146  to the trailing edge  130 . 
     The streamline  150  can flow over the dimples  126  of the airfoil  100  such that a portion of the streamline  150  can flow into the dimples  126 . The portion of the streamline  150  that flows into the dimples  126  can follow the contour of the dimples  126 , specifically the contour the non-perforated section  136  of the dimples  126 , as illustrated. As the streamline  150  flows from the leading edge  128  to the maximum depth  146  of a corresponding dimple  126 , the portion of the streamline  150  within the dimple  126  can be accelerated. This is due to the fact that the at least a portion of the first and second lateral walls  142 ,  144  converge from the leading edge  128  to the trailing edge  130  of the dimples  126  such that the portion of the streamline  150  within the dimples  126  is compressed and subsequently accelerated within the dimples  126 . As the portion of the streamline  150  within the dimples reaches the maximum depth  146 , it is redirected upwards or towards the trailing edge  130  of the dimples  126  by the second bottom wall  140  of the non-perforated section  136 . This flow of the portion of the streamline  150  within the dimples  126  from the leading edge  128 , over the first bottom wall  138 , to the maximum depth  146 , and up the second bottom wall  140  to the trailing edge  130  can define vortices  156  within the dimples  126  generated by the portion of the streamline  150  that flows through the dimples  126 . The vortices  156  can be defined as Kelvin Helmholtz vortices, which directly oppose the streamline  150  flowing over the dimples  126 . 
     In other words, the dimples  126  can be shaped to generate the vortex  156  that rotates counter or otherwise opposes to the streamline  150  flowing over the airfoil  100 , and specifically the dimples  126 , in the stream-wise direction. These vortices  156  can be configured to counteract the streamline  150  such that the streamline  150  is suctioned to the outer wall  110  of the airfoil  100 . 
     It is contemplated that the boundary layer growth  154  can occur downstream of where the boundary layer growth  254  would occur. This is due to the fact that the airfoil  200  does not include the set of dimples  126  while the airfoil  100  does. As such, the vortices  156  are only generated by the airfoil  100  and not by the exemplary airfoil  200 . As the vortices  156  of the airfoil  100  hold or suction the streamline  150  to the outer wall  110  of the airfoil  100 , the streamline  150  does not separate from the outer wall  110  to form the boundary layer growth  154  until farther downstream than where the boundary layer growth  254  occurs. The sooner that the boundary layer growth  154 ,  254  occurs, the larger the total boundary layer growth  154 ,  254 . The more boundary layer growth  154 ,  254  there is, the more pressure losses occur, which results in a decreased power output or torque generated by the airfoils and an overall decrease in the efficiency of the gas turbine engine  10 . As the boundary layer growth  154  occurs further downstream the airfoil  100  than the boundary layer growth  254 , the boundary layer growth  154  is ultimately smaller than the boundary layer growth  254 . As such, the pressure losses from the leading edge  116  to the trailing edge  118  of the airfoil  100  are smaller than the pressure losses from the leading edge  116  to the trailing edge  118  of the airfoil  200 . This ultimately results in the torque generated from the airfoil  100  being larger than the torque generated by the airfoil  200 , which ultimately results in the gas turbine engine  10  including the airfoil  100  with the set of dimples  126  being more efficient than an exemplary gas turbine engine that does not include the airfoil  100  and only the airfoil  200 . 
       FIG. 7  is a planform view of an exemplary dimple  326  of the set of dimples  126  of  FIG. 2 . The dimple  326  is similar to the dimple  126 ; therefore, like parts will be identified with like numerals in the  300  series, with it being understood that the description of the like parts of the dimple  126  applies to the dimple  326  unless otherwise noted. 
     The dimple  326  can be demarcated by the leading edge  128 , a trailing edge  130 , a first lateral edge  332  and a second lateral edge  334 . A non-perforated section  336  of the dimple  326  can include a first bottom wall  338  extending from the leading edge  128  to a maximum depth  346  and a second bottom wall  340  extending from the maximum depth  346  to the trailing edge  330 . As illustrated the first and second lateral edges  332 ,  334  can converge to an intersection point defined by the trailing edge  330 . As such, when viewed in planform, the non-perforated section  336  can define a triangular two-dimensional surface area that decreases from the leading edge  128  to a trailing edge  330  of the dimple  326 . 
     The non-perforated section  336 , similar to the non-perforated section  136 , can decrease in depth from the leading edge  128  to the maximum depth  346 , and increase from the maximum depth  346  toward the trailing edge  330  or to the first or second lateral edges  332 ,  334 . As illustrated, the maximum depth  346  is defined as a point. As such, the non-perforated section  336 , defined by the second bottom wall  340 , can non-continuously or non-linearly decrease in depth from the point at which the maximum depth  346  occurs to the at least a portion of the first and second lateral edges  332 ,  334  proximate the trailing edge  330 . Alternatively, the maximum depth can be formed as two lines converging from the first and second lateral edges  332 ,  334  toward a vertex. In such a case, the second bottom wall  340  can non-continuously, or non-linearly decrease in depth to the trailing edge  330 . It will be appreciated that the maximum depth  346  can occur at any point along a span-wise length of the dimple  326 . For example, the maximum depth  346  can be equidistant from between the first and second lateral edges  332 ,  334 . Alternatively, the maximum depth  346  can be closer to one of either the first or second lateral edges  332 ,  334  than the other. 
       FIG. 8  is a planform view of another exemplary dimple  426  of the set of dimples  126  of  FIG. 2 . The dimple  426  is similar to the dimple  126 ,  326 ; therefore, like parts will be identified with like numerals in the  400  series, with it being understood that the description of the like parts of the dimple  126 ,  326  applies to the dimple  426  unless otherwise noted. 
     The dimple  426  can include a leading edge  428 , a trailing edge  430 , a first lateral edge  432 , and a second lateral edge  434 . A non-perforated section  436  of the dimple  426  can include a first bottom wall  438  extending from the leading edge  428  to a maximum depth  446  and a second bottom wall  440  extending from the maximum depth  446  to the trailing edge  430 . The leading edge  428  and the trailing edge  430  can each include a filleted or rounded edge  458 . Specifically, the leading edge  428  can include two rounded edges  458 , each opposing each other at opposites span-wise ends of the leading edge  428 , that join or merge with a first lateral edge  432  and a second lateral edge  434 . The first and second lateral edges  432 ,  434  can converge toward the rounded edges  458  of the trailing edge  430 . As such, when viewed in planform, the non-perforated section  436  of the dimple  426  can define a rounded vertex, triangular two-dimensional surface area that decreases from the leading edge  428  to a trailing edge  430  of the dimple  426 . 
       FIG. 9  is a planform view of another exemplary dimple  526  of the set of dimples  126  of  FIG. 2 . The dimple  526  is similar to the dimple  126 ,  326 ,  426 ; therefore, like parts will be identified with like numerals in the  500  series, with it being understood that the description of the like parts of the dimple  126 ,  326 ,  426  applies to the dimple  526  unless otherwise noted. 
     The dimple  526  can include the leading edge  128 , a trailing edge  530 , the first lateral edge  132 , and a second lateral edge  534 . A non-perforated section  536  of the dimple  526  can include a first bottom wall  538  extending from the leading edge  128  to a maximum depth  546  and a second bottom wall  540  extending from the maximum depth  546  to the trailing edge  530 . As illustrated, the first and second lateral edges  132 ,  534  can extend between the leading edge  128  and the trailing edge  530  at differing rates. Specifically, the first lateral edge  132  can be angled to converge from the leading edge  128  toward the trailing edge  530 , while the second lateral edge  534  can be normal to both the leading and trailing edges  128 ,  530 . As such, when viewed in planform, the non-perforated section  536  of the dimple  526  can define an exemplary two-dimensional surface area, specifically a non-symmetric trapezoidal two-dimensional surface area, that decreases from the leading edge  128  to the trailing edge  530  of the dimple  526 . 
       FIG. 10  is a planform view of another exemplary dimple  626  of the set of dimples  126  of  FIG. 2 . The dimple  626  is similar to the dimple  126 ,  326 ,  426 ,  526 ; therefore, like parts will be identified with like numerals in the  600  series, with it being understood that the description of the like parts of the dimple  126 ,  326 ,  426 ,  526  applies to the dimple  626  unless otherwise noted. 
     The dimple  626  can include the leading edge  128 , a trailing edge  630 , a first lateral edge  632 , and a second lateral edge  634 . A non-perforated section  636  of the dimple  626  can include a first bottom wall  638  extending from the leading edge  128  to a maximum depth  646  and a second bottom wall  640  extending from the maximum depth  646  to the trailing edge  630 . As illustrated, the first and second lateral edges  632 ,  634  can be defined as a set of curved edges that converge from leading edge  128  toward the trailing edge  630 , with the first and second lateral edges  632 ,  634  intersecting at the trailing edge  630 . As such, when viewed in planform, the non-perforated section  636  of the dimple  626  can define a semi-oval two-dimensional surface area that decreases from the leading edge  128  to the trailing edge  630  of the dimple  626 . 
     Benefits of the present disclosure include an increased efficiency of the gas turbine engine when compared to conventional gas turbine engines. For example, conventional gas turbine engines can include at least one engine components exposed to a fluid flow and shaped like an airfoil (e.g., airfoils in the compressor section or turbine section, fan blades, guide vanes, etc.). The fluid flow can flow over the engine components from a leading edge toward a trailing edge following a contour of an outer wall on a suction side of the engine component to define a streamline. The streamline will separate from the outer wall to define the boundary layer growth. The sooner that the boundary layer growth begins to form, the larger the total boundary layer growth there will be. This can ultimately negatively affect overall efficiency of the gas turbine engine by increasing pressure losses associated with the flow of fluid over the engine component and reducing the torque generated by the engine component. The gas turbine engine as described herein, however, include a set of airfoils including the set of dimples provided along a portion of the outer wall on the suction side of the airfoils. As discussed herein, the shape of the dimples can generate vortices that that counter-act the streamline. Specifically, the vortices can be generated by the dimples including non-perforated section, which increases in depth from the leading edge of the dimples to the maximum depth that is closer to the trailing edge than the leading edge, and then decreases in depth from the maximum depth to the trailing edge, along with the convergence of at least one of the first or second lateral edges or walls from the leading edge to the trailing edge. The vortices ensure the streamline will follow a contour defined by the outer wall for a larger stream-wise portion of the airfoil, or other engine component, when compared to the engine components in conventional gas turbine engines. As such, the pressure losses associated with the boundary layer growth are smaller in the gas turbine engine as described herein when compared to conventional gas turbine engines, and a greater torque is generated by the airfoil or engine component including the set of dimples when compared to conventional engine components. With a decrease in the boundary layer growth, a decrease in pressure losses, and an increase in torque, the overall efficiency of the gas turbine engine including the airfoils with the dimples is ultimately increased when compared to traditional gas turbine engines. 
     To the extent not already described, the different features and structures of the various aspects can be used in combination with each other as desired. That one feature cannot be illustrated in all of the aspects is not meant to be construed that it cannot be, but is done for brevity of description. Thus, the various features of the different aspects can be mixed and matched as desired to form new aspects, whether or not the new aspects are expressly described. Combinations or permutations of features described herein are covered by this disclosure. 
     This written description uses examples to describe aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of aspects of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims. 
     Further aspects of the invention are provided by the subject matter of the following clauses: 
     A gas turbine engine comprising a set of circumferentially adjacent airfoils, the airfoils having an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction, and a set of dimples provided on the outer wall of at least one of the airfoils, with the dimples spaced in at least one of the stream-wise and or span-wise direction, and the dimples having a non-perforated section in a shape defining a two-dimensional surface area that when viewed in planform, the two-dimensional surface area decreases in the stream-wise direction. 
     The gas turbine engine of any preceding clause, wherein the dimples are arranged in rows in the stream-wise direction, and at least one row is staggered with respect to the other rows. 
     The gas turbine engine of any preceding clause, wherein the dimples are arranged in columns in the span-wise direction. 
     The gas turbine engine of any preceding clause, wherein the non-perforated section increases in depth in the stream-wise direction. 
     The gas turbine engine of any preceding clause, wherein the dimples have a leading edge and a trailing edge to define a stream-wise length between the leading edge and trailing edge, and the depth first increases from the leading edge to a maximum depth and then the depth then deceases from the maximum depth to the trailing edge. 
     The gas turbine engine of any preceding clause, wherein the maximum depth is between greater than 50% and less than 100% of the stream-wise length. 
     The gas turbine engine of any preceding clause, wherein a portion of the non-perforated section from the maximum depth to the trailing edge defines a curved cross section when viewed in the span-wise direction. 
     The gas turbine engine of any preceding clause, wherein the depth constantly increases from the leading edge to the maximum depth. 
     The gas turbine engine of any preceding clause, wherein the depth constantly decreases from the maximum depth to the trailing edge. 
     The gas turbine engine of any preceding clause, wherein the airfoils are provided within at least a portion of at least one of a compressor section, a turbine section, a fan section, an exhaust section or a bypass duct of the gas turbine engine. 
     The gas turbine engine of any preceding clause, wherein the airfoils are provided within at least a portion of a low-pressure turbine of the turbine section. 
     The gas turbine engine of any preceding clause, wherein the dimples are shaped to generate a vortex that rotates counter to a streamline flowing over the at least one of the airfoils in the stream-wise direction. 
     An airfoil comprising an outer wall defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a stream-wise direction, and between a root and a tip to define a span-wise direction, and a set of dimples provided on the outer wall, with the dimples spaced in at least one of the stream-wise or span-wise direction, and the dimples having a non-perforated section in a shape defining a two-dimensional surface area that when viewed in planform, the two-dimensional surface area decreases in the stream-wise direction. 
     The airfoil of any preceding clause, wherein the dimples are arranged in rows in the stream-wise direction, and at least one row is staggered with respect to the other rows. 
     The airfoil of any preceding clause, wherein the dimples are arranged in columns in the span-wise direction. 
     The airfoil of any preceding clause, wherein the non-perforated section constantly increases in depth in the stream-wise direction. 
     The airfoil of any preceding clause, wherein the dimples have a leading edge and a trailing edge to define a stream-wise length between the leading edge and trailing edge, and the depth first increases from the leading to a maximum depth and then the depth then deceases from the maximum depth to the trailing edge. 
     The airfoil of any preceding clause, wherein the maximum depth is between greater than 50% and less than 100% of the stream-wise length. 
     The airfoil of any preceding clause, wherein a portion of the non-perforated section from the maximum depth to the trailing edge defines a curved cross section when viewed in the span-wise direction. 
     The airfoil of any preceding clause, wherein the dimples are shaped to generate a vortex that rotates counter to a streamline flowing over the airfoil in the stream-wise direction.