Patent Publication Number: US-2020277862-A1

Title: Airfoil for a turbine engine

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation application of U.S. patent application Ser. No. 15/138,624, filed Apr. 26, 2016, which is incorporated herein by reference in its entirety. 
    
    
     TECHNICAL FIELD 
     This disclosure relates generally to a turbine engine, and more specifically to a method of forming at least a portion of a wall of a component of the turbine engine. 
     BACKGROUND OF THE INVENTION 
     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 rotating turbine blades. 
     Turbine engines for aircraft are often designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components can be beneficial or necessary. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components that require cooling. Temperatures in the high pressure turbine can be 1000° C. to 2000° C. and the cooling air from the compressor can be around 500° C. to 700° C. While the compressor air is a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine. 
     Contemporary engine components, such as the rotating blades, necessarily account for a portion of the overall engine weight. Decreasing the weight of these engine components is desirable to increase engine efficiency. Decreasing weight of the engine components can be accomplished by utilizing thinner walls for the components, for example. However, thinner walls include a decreased volume through which film holes can extend, which can decrease the effectiveness of the film holes. Thus, it is desirable to utilize thinner walls for the engine components to decrease system weight while providing sufficient length for the film holes to maintain cooling effectiveness. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, embodiments of the invention relate to a component for a turbine engine, with the turbine engine generating a hot combustion gas flow and providing a cooling fluid flow, the component comprising a wall having a nominal thickness separating the hot combustion gas flow from the cooling fluid flow having a hot surface facing the hot combustion gas flow and a cool surface facing the cooling fluid flow, at least one localized, protuberance extending from the cool surface, and a film hole extending through the protuberance and the wall, having a length greater than the nominal thickness of the wall. 
     In another aspect, embodiments of the invention relate to an airfoil for a turbine engine comprising a wall having a first side adjacent to a first fluid flow and a second side adjacent to a second fluid flow and having a nominal thickness, at least one localized, protuberance extending from the wall, and a cooling hole extending through the protuberance and the wall, having a length greater than the nominal thickness of the wall. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine for an aircraft. 
         FIG. 2  is an isometric view of an airfoil of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a cross-sectional view of the airfoil of  FIG. 2  having walls including protuberances with film holes. 
         FIG. 4  is a perspective view of the wall having the protuberance with the film hole extending through the protuberance and the wall. 
         FIG. 5  is a cross-sectional view of the wall of  FIG. 4  illustrating a profile of the protuberance and film hole across a direction of a cooling fluid flow. 
         FIG. 6  is a cross-sectional view of the wall of  FIG. 4  illustrating a profile of the protuberance and film hole in the direction of the cooling fluid flow. 
         FIG. 7  is a cross-sectional view of another embodiment of the film hole extending in a direction opposite of the cooling fluid flow. 
         FIG. 8  is a cross-sectional view of yet another embodiment the film hole having an inlet on a forward face of the protuberance. 
         FIG. 9  is a cross-sectional view of the protuberance having a recess with the film hole inlet disposed in the recess. 
         FIG. 10  is a perspective view of an elongated protuberance having a film hole offset form the direction of the cooling fluid flow. 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     The described embodiments of the present invention are directed to an engine component for a turbine engine having at least one protuberance disposed on a wall of the engine component with a film hole extending through the protuberance and the wall. For purposes of illustration, the present invention will be described with respect to the turbine for an aircraft gas turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability within an engine, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
     As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component. 
     Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. 
     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 invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. 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 engine  10  has a generally longitudinally extending axis or centerline  12  extending forward  14  to aft  16 . The 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 plurality of fan blades  42  disposed radially about the centerline  12 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10 , which generates combustion gases. The 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 centerline  12  of the 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 centerline  12  of the 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 and couple to a plurality of rotatable elements, which can collectively define a rotor  51 . 
     The LP compressor  24  and the HP compressor  26  respectively include a plurality 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 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 plurality 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 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 engine  10 , such as the static vanes  60 ,  62 ,  72 ,  74  among the compressor and turbine section  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 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 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 draw 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 engine assembly  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a plurality 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 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 . 
       FIG. 2  is a perspective view of an engine component in the form the turbine blades  68  of the engine  10  of  FIG. 1 . It should be understood that the turbine blade  68  is exemplary and that the engine component can include other components requiring cooling. The turbine blade  68  includes a dovetail  90  and an airfoil  92 . The airfoil  92  includes a tip  94  to a root  96 . A span-wise direction can be defined between the tip  94  and the root  96 . The dovetail  90  includes a platform  98  and one or more inlet passages  100  having an outlet  102 . The dovetail  90  and platform  98  can be integral with the airfoil  92  adjoining at the root  96 . The platform  98  helps to radially contain the turbine airflow driven by the airfoil  92 . The dovetail  90  can be configured to mount to a turbine rotor disk  71  ( FIG. 1 ) to rotate the airfoil  92  about the engine centerline  10 . The inlet passages  100  can be fed with a flow of air, such as bypass air  104 . The bypass air  104  is provided to the airfoil  92  at the root  96  exhausting through the outlets  102 . It should be appreciated that the dovetail  90  is shown in cross-section, such that the inlet passages  100  are housed within the body of the dovetail  90 . 
     Referring to  FIG. 3 , the airfoil  92 , shown in cross-section, has an interior  110  bounded by an outer wall  112 . The outer wall  112  includes a concave-shaped pressure sidewall  114  and a convex-shaped suction sidewall  116 . A leading edge  118  and a trailing edge  120  are defined at the junction between the pressure and suction sidewalls  114 ,  116 , defining a chord-wise distance between the leading and trailing edges  118 ,  120 . The airfoil  92 , when implemented as a rotating blade as compared to a stationary vane, rotates in a direction such that the pressure sidewall  114  follows the suction sidewall  116 . Thus, as shown in  FIG. 3 , the airfoil  92  would rotate upward toward the top of the page. 
     One or more ribs  130  are included in the interior  110 . The ribs  130  can extend between the pressure and suction sidewalls  114 ,  116  to define internal chambers  132 . The chambers  132  can be discrete compartments defined within the airfoil  92 . Alternatively, the chambers  132  can be in fluid communication with one another, such as defining a serpentine flow path snaking through the airfoil  92  in the span-wise direction. It should be understood that the ribs  130  and chambers  132  defined by the ribs  130  are exemplary and should not be construed as limiting. It is contemplated that the interior  110  or chambers  132  defined therein can also include a plurality of plenums, circuits, micro-circuits, near wall cooling systems, pin banks, or similar structures in non-limiting examples. 
     A protuberance  134  can be disposed on the outer wall  112 . The protuberances  134  are discrete members, defining an increased thickness for the wall  112 . In non-limiting examples, the protuberance can be radiused, rounded, conical, frustoconical, bell-shaped, or non-linear. Additional examples of protuberances can include, but are not limited to, radiused, circular, oval, elliptical, spherical, ellipsoidal, or curvilinear. The protuberances  134  can be integral to the outer wall  112 , or can be mounted thereto. In one non-limiting example, the protuberances  134  can be formed on the airfoil  92  using additive manufacturing. Any number of protuberances  134  can be included on the outer wall  112  or the ribs  130  and can be organized in any manner, such as a linear grouping in the span-wise or chord-wise direction, a pattern, or random placement. 
     Additionally, it is contemplated that the protuberances  134  can be formed on an interior wall of the airfoil  92  or an engine component. In one example, the protuberances  134  can be formed on the rib  130 . In other non-limiting examples, the protuberance  134  can be disposed on walls of cooling structures, such as micro-circuits, cooling mesh, plenums, pin banks, or other component structures requiring cooling. 
     A film hole  136  can extend through the protuberances  134 . The film holes  136  can extend through the protuberances  134  on the outer wall  112  to provide a cooling film along the external surface of the outer wall  112  for cooling the airfoil  92 . Additionally, in the case where the protuberance  134  is disposed on an interior wall or structure, such as the rib  130  in one non-limiting example, the film hole  136  can be a cooling hole such as a cross-over hole. In additional examples, such a cooling hole can provide a flow of cooling fluid among internal cavities or chambers of the engine component, such as adjacent areas channels or a micro circuit. 
     The film hole  136  can be shaped to direct a flow of fluid entering the film hole  132 , passing through the film hole  132 , or exhausting from the film hole  132 . Such shaping, for example, can include a converging, diverging, or metering section to direct the flow of fluid, in non-limiting examples. The converging section can increase the flow velocity of the flow of fluid, the diverging section can decrease the flow velocity of the flow of fluid, and the metering section can meter the flow of fluid passing through it. Additional shaping can include an expansion section or a reduction section. The expansion section can include an increasing cross-sectional area to form a diffusion section and the reduction section can include a decreasing cross-sectional area. Additionally, the shaping of the film hole  136  can include a non-linear film hole  136 . Such a film hole  136  could include curved passages or follow the curvature of the protuberance  134 . 
     It should be understood that the protuberance  134  could be placed on any wall having opposing flows on opposing sides of the wall  118 . Additionally, the film hole  136  can pass through the protuberance to provide a flow between the opposing sides of the wall  118 . 
     In an example where the engine component is not a blade, such as a vane, combustion liner, shroud, or other component requiring cooling, the protuberances can be disposed an any wall, such as an internal or external wall, and can include a film hole to provide a flow of fluid through such a wall for providing a cooling film. Thus, it should be understood that the airfoil  92  as illustrated is exemplary and non-limiting, and the protuberances  134  can have equal applicability in any other engine component utilizing film holes. 
       FIG. 4  illustrates one protuberance  134  having a film hole  136  extending through the protuberance  134 . The protuberance can be non-rectilinear, including a non-linear surface extending from the first side  144  to the inlet  150 , with the inlet  150  being rounded, transitioning into the film hole  136 . In the top view shown, the protuberance is circular. In other non-limiting examples, the protuberance can be oval, elliptical, spherical, ellipsoidal, or curvilinear. The protuberance  134  can be symmetrical, being even about an axis parallel to the direction of a cooling fluid flow C. The protuberance  134  is disposed on a wall  140 . The wall  140  can be the outer wall  112 , for example, or any other component wall having a film hole  136 . The wall  140  has a nominal thickness  142 , having a first side  144  and a second side  146  defining a consistent nominal thickness  142  between the sides  144 ,  146 . The protuberance  134  is a circular extension, extending from the first side  144  into a first fluid flow  148 . The film hole  136  is disposed in the center of the circular protuberance  134 , having an inlet  150  and an outlet  152 . A passage  154  couples the inlet  150  to the outlet  152 . 
     The nominal thickness  142  can be the thickness for the wall  140  defined as the distance between the first and second side  144 ,  146 . Such a nominal thickness  142  can be determined in many different ways. For example, the nominal thickness  142  for the wall can be a function of the thermal load on the wall  140 , the airfoil  92 , or the engine component. In other examples, the nominal thickness  142  can be a function of a vibratory force acting on the wall  140 , a pressure differential between on opposing sides of the wall  140 , or the manufacturer required load for the wall  140  during operation. It should be appreciated that the nominal thickness  142  can be determined by multiple methods, such that a minimum operational thickness for the particular wall  140  is determined. Additionally, the nominal thickness  142  is a minimal operational thickness of the wall  140 , being a function of the thermal load, vibratory force, pressure differentials, load requirements, or other similar method can be respective of minimal operation requirements to maintain safe operation of the engine and individual engine components. The nominal thickness  142  for the wall  140  can reduce engine weight, increasing engine efficiency or performance. 
     Referring to  FIG. 5 , the first side  144  can be a cool side or cool surface, adjacent to a first fluid flow  148 , such as a cooling fluid flow C. The second side  146  can be a hot side, or a hot surface, adjacent to a second fluid flow  170 , such as a hot gas flow H. The profile view of the protuberance  134  in  FIG. 5  illustrates a height  160  of the protuberance  134 . The height  160  is the maximum distance the protuberance  134  extends from the wall  140 . The height  160  can be determined in multiple different ways. For example, the height  160  can be a function of the nominal thickness  142 . In one example, the height  160  can be at least 50% of the nominal thickness  142 . In another example, the height  160  can be equal to the nominal thickness  142 , or greater. In yet another example, the height can be equal to at least 100% of the nominal thickness  142 . It should be appreciated that the height  160  can be anywhere from 5% of the nominal thickness  142  to 200% of the nominal thickness  142  or greater. 
     Alternatively, the height  160  can be a function of the film hole  136 . A length  162  of the film hole  136  can be defined as the distance between the inlet  150  and the outlet  152 . The height  160  can be a function of the length  162 , where a particular film hole  136  can require a particular length  162  to provide an effective flow of fluid. For example, the height  160  can be at least 50% of the length  162 . While  FIG. 5  illustrates a linear film hole  136 , it should be understood that the film hole  136  need not be linear, and with such a film hole, the length  162  can be measured as the streamline distance between the inlet  150  and the outlet. It should be appreciated that the film hole  136  as illustrated in  FIG. 5  is a perpendicular film hole  136 . In the case of the non-linear film hole, the length  162  will increase. As such, it should be appreciated that in such a case, the height  160  can be at least 30% of the length  162 . 
     In yet another example, the height  160  can be determined as a function of a diameter  164  of the film hole  136 . A particular diameter  164  for a film hole may be required by an engine component, in order to keep structural integrity of the engine component. The diameter  164  can require a particular length  162  for the film hole  136  to maintain an effectiveness, defining a required length-to-diameter ratio (L/D) for the film hole  136 . As such, the diameter  164 , or the L/D ratio can be used to define the height  160  in order to provide sufficient film hole effectiveness. 
     Referring now to  FIG. 6 , a side profile view of the protuberance  134  illustrates one orientation of the film hole  136 . The protuberance  134  can be conical, having a portion removed at the inlet  150  of the film hole  136 . Alternatively, it is contemplated that the protuberance  135  can have a conic profile, having the inlet  150  disposed on one of the sides of the protuberance. The first cooling flow  148  can be a cooling fluid flow C. A second fluid flow  170 , adjacent to the second side  146  of the wall  140 , can be a hot gas flow H. The film hole  136  can be angled in the direction of one of the first and second fluid flows  148 ,  170 , or both. As such, the inlet  150  can be positioned upstream of the outlet  152  relative to the cooling fluid flow C. In an example where the film hole  136  is a film hole, the cooling fluid flow C can be provided through the film hole  136  to the second side  146  as a cooling film  172  to cool the engine component. 
       FIG. 7  illustrates another embodiment of the protuberance  134  having a rounded dimension with a frustoconical shape at the inlet  150  of the film hole  136 . The film hole  136  includes the inlet  150  disposed downstream of the outlet  152  relative to the cooling fluid flow C. Such an orientation can be advantageous for providing an effective film hole length as well as providing multiple directional capabilities for exhausting a fluid from the film hole  136 .  FIG. 8  illustrates another embodiment, having a rounded protuberance having the film hole  136  with an inlet  150  offset from the center of the protuberance  134 . The protuberance  134  can be divided into an upstream side  174  and a downstream side  176 . The inlet  150  can be disposed on the upstream side  174 . Such an orientation can be advantageous for determining flow rate entering the film hole  136 . Alternatively, it is contemplated that the inlet  150  can be disposed on the downstream side  176 , or any other position on the protuberance  134 . 
     It should be appreciated that the position and orientation of the film hole  136  of  FIGS. 5-8  is exemplary. The position of the inlet  150 , outlet  152 , and dimension of the passage  154  disposed therebetween can be adapted to control flow rates through the film hole  136 , or adapt the length  162 , diameter  164 , or length-to-diameter ratio for the film hole  136  to provide effective cooling through the film hole  136 . Furthermore, it is contemplated that the film hole  136  can be provided with inlet shaping or outlet shaping, to provide a more deterministic flow for a cooling fluid passing through the cooling flow. Such an example would be a diverging outlet which can provide a cooling fluid over a greater cross-sectional area of the engine component. 
     Additionally, it should be appreciated that the protuberance  134  can have a height  160  dependent on portions of the engine component, such as the nominal thickness  142 , the length  162 , diameter  164 , or L/D ratio of the film hole  136 . 
     Referring now to  FIG. 9 , a recess  180  can be formed in the protuberance  134 . The recess  180  can be machined as part of the protuberance  134 , such as during additive manufacturing, or can be removed from the protuberance  134  to form the recess  180 . The recess  180  can be symmetrical, such as a hemispherical shape, while any shape is contemplated. In other non-limiting examples, the recess  180  can be a rectilinear shape, or an arcuate or radiused shape, or any combination thereof. The film hole  136  can be disposed in the recess  180 , having the inlet  150  at least partially formed within the recess  180 . The size or shape of the recess  180  can be used to control the flow rate of a flow of fluid provided to the film hole  136 , or to further reduce component weight in combination with the nominal thickness  142  for the wall  140 . 
     Referring now to  FIG. 10 , the protuberance  134  can be asymmetrical, having an elongated or offset shape. Such as shape may be desirable to optimize fluid flows within the engine component or for directing a flow toward one or more film holes  136 . Additionally the film hole  136  can be offset from the direction of the first fluid flow  148  along the engine component. The discrete direction of the first fluid flow  148  at the protuberance  134  can be transposed on the protuberance  134  as a transposed axis  190 . The passage  154 , shown in a linear example while non-linear film holes are contemplated, can define a passage axis  192 . A film hole angle  194  can be defined between the transposed axis  190  and the passage axis  192  to define the offset relationship of the film hole  136  to the first fluid flow  148 . Additionally, the film hole  136  can be angled relative to a local normal between the first and second sides  144 ,  146 . 
     It should be understood that the offset orientation of the film hole  136  or the protuberance  134  can be discrete, relative to an adjacent flow of fluid which can change direction or magnitude at different portions of the engine component. As such, a plurality of protuberances  134  along the engine component can be aligned or patterned, while some of the protuberances  134  or film holes  136  are offset from the direction of the first fluid flow  148  at a portion of the engine component. 
     A method of cooling an engine component, such as the airfoil  92 , can include a cool surface, such as the first side  144 . The method can include passing a cooling fluid flow C along the cool surface  144  and providing at least a portion of the cooling fluid flow C through a film hole, such as the film hole  136 , in a protuberance  134  extending from the cool surface  144 . Providing at least a portion of the cooling fluid flow C can include providing a portion of the cooling fluid flow through a recess  180  in the protuberance  134  prior to providing the cool fluid flow C to the film hole  136 . Additionally, providing a portion of the cooling fluid flow C through the recess  180  can minimize dust accumulation at the film hole  136  or along the cool surface  144  of the engine component. 
     It should be understood that the airfoil  92  or other engine component requiring cooling can utilizing the film hole  136  such as a film hole disposed within the protuberance  134 . The protuberance  134  provides for an increased thickness permitting an increased film hole length  162  to provide effective cooling through the film hole  136 . At the same time, the use of a protuberance  134  permits the remaining portions of the engine component to have a nominal thickness  142 , which reduces component weight, reducing overall engine weight. A reduced weight provides for better engine efficiency. 
     It should also be understood that the protuberances  134  are discrete, having no greater an area than necessary to provide for the casting, drilling, or otherwise forming the film holes  136  through the protuberances  134  in order to have an increased length, diameter, or L/D ratio for the film hole  136  which would otherwise be unachievable within the nominal thickness  142  of the engine component, due to the nominal thickness  142  to manufacturing capabilities of the engine component at the nominal thickness  142 . 
     Furthermore, the protuberances  134  are radiused, reducing drag or resistance caused by the extension of the protuberance  134  into the flow of fluid, such as the cooling fluid flow C, adjacent the protuberance  134 . Further still, the radiused protuberances  134  or recesses  180  therein can provide for reduced dust accumulation, increasing component lifetime or reducing required maintenance. 
     It should be appreciated that the airfoil, engine components, protuberances, or film holes described herein can be formed by additive manufacturing. Such manufacturing can be used to develop the intricate details of the aforementioned, such as specific film hole shaping without the poor yields of such manufacturing as casting, or the imperfections associated with other manufacturing methods such as film hole drilling. 
     It should be appreciated that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may 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.