Patent Publication Number: US-2023145370-A1

Title: Turbomachine blade trailing edge cooling circuit with turn passage having set of obstructions

Description:
TECHNICAL FIELD 
     The disclosure relates generally to turbomachine blades and, more particularly, to a turbomachine blade having a trailing edge cooling circuit with a turn passage having a set of obstructions therein. 
     BACKGROUND 
     Turbomachine blades, such as rotor blades or stationary vanes, include airfoils that accelerate flow through contraction of area and the introduction of tangential velocity. The trailing edges of the airfoils are difficult to cool due to the small volume of material compared to the large heat loads at that location. Notably, the mismatch between external surface area and the internal surface makes any cooling solution challenging. To address this situation, trailing edges are typically cooled with coolant flows having high flow rates. The high flow rates to the trailing edges decreases the coolant that can be used elsewhere. The high flow rates also require the trailing edges to have minimum thicknesses to accommodate the passages that deliver the coolant flow and create the necessary cold-to-hot area ratio. The minimum thicknesses do not allow for sharper trailing edges that would improve aerodynamic performance. 
     BRIEF DESCRIPTION 
     All aspects, examples and features mentioned below can be combined in any technically possible way. 
     An aspect of the disclosure provides a turbomachine blade, comprising: an airfoil body having a pressure side and a suction side connected by a leading edge and a trailing edge; a coolant feed passage defined in the airfoil body; a first coolant reuse passage defined in the airfoil body; a first cooling circuit defined in the airfoil body, the first cooling circuit including: a first rearward passage extending toward the trailing edge from and fluidly coupled to the coolant feed passage; a first radially spreading return passage extending away from the trailing edge toward and fluidly coupled to the first coolant reuse passage; and a first radially extending turn passage coupling the first rearward passage and the first radially spreading return passage; and a first set of obstructions positioned in the first radially extending turn passage. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a plurality of vent passages extending from the first radially extending turn passage through the trailing edge of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first rearward passage is radially offset from the first radially spreading return passage along a radial axis of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a second set of obstructions positioned in the first radially spreading return passage. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises: a second cooling circuit defined in the airfoil body, the second cooling circuit including: a second rearward passage extending toward the trailing edge from and fluidly coupled to the coolant feed passage; a second radially spreading return passage extending away from the trailing edge toward and fluidly coupled to a second coolant reuse passage defined in the airfoil body; and a second radially extending turn passage coupling the second rearward passage and the second radially spreading return passage; wherein the second rearward passage is radially offset from the second radially spreading return passage along the radial axis of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first cooling circuit is circumferentially offset from the second cooling circuit in the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a second set of obstructions positioned in the second radially spreading return passage. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a plurality of the first cooling circuits radially spaced in the airfoil body, and a plurality of second cooling circuits radially spaced in the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the trailing edge has an ellipse ratio between 1.1 and 4. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first cooling circuit is adjacent the suction side of the airfoil body, and the second cooling circuit is adjacent the pressure side of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first radially extending turn passage has a first circumferential width at a forward end thereof that is greater than a second circumferential width at an aft end thereof. 
     An aspect of the disclosure provides a coupon for replacing a cutout of a predetermined area in an airfoil body of a turbomachine blade, the airfoil body having a pressure side and a suction side connected by a leading edge and a trailing edge, the cutout within the trailing edge of the airfoil body, the coupon comprising: a coupon body; a first cooling circuit defined in the coupon body, the first cooling circuit including: a first rearward passage extending toward the trailing edge from and fluidly coupled to a coolant feed passage defined in at least one of the coupon body and the airfoil body; a first radially spreading return passage extending away from the trailing edge toward and fluidly coupled to a first coolant reuse passage defined in at least one of the coupon body and the airfoil body; a first radially extending turn passage coupling the first rearward passage and the first radially spreading return passage; and a first set of obstructions positioned in the first radially extending turn passage. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a plurality of vent passages extending from the first radially extending turn passage through the trailing edge of the coupon body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first rearward passage is radially offset from the first radially spreading return passage along a radial axis of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises a second set of obstructions positioned in the first radially spreading return passage. 
     Another aspect of the disclosure includes any of the preceding aspects, and further comprises: a second cooling circuit defined in the coupon body, the second cooling circuit including: a second rearward passage extending toward the trailing edge from and fluidly coupled to the coolant feed passage; a second radially spreading return passage extending away from the trailing edge toward and fluidly coupled to a second coolant reuse passage defined in at least one of the coupon body and the airfoil body; and a second radially extending turn passage coupling the second rearward passage and the second radially spreading return passage, wherein the second rearward passage is radially offset from the second radially spreading return passage along the radial axis of the airfoil body. 
     Another aspect of the disclosure includes any of the preceding aspects, and the trailing edge has an ellipse ratio between 1.1 and 4. 
     Another aspect of the disclosure includes any of the preceding aspects, and the first radially extending turn passage has a first circumferential width at a forward end thereof that is greater than a second circumferential width at an aft end thereof. 
     An aspect of the disclosure provides a gas turbine system, comprising: a compressor; a combustor; and a turbine, the turbine including a turbomachine blade including a trailing edge cooling system, the turbomachine blade including: an airfoil body having a pressure side and a suction side connected by a leading edge and a trailing edge; a coolant feed passage defined in the airfoil body; a first coolant reuse passage defined in the airfoil body; a first cooling circuit defined in the airfoil body, the first cooling circuit including: a first rearward passage extending toward the trailing edge from and fluidly coupled to the coolant feed passage; a first radially spreading return passage extending away from the trailing edge toward and fluidly coupled to the first coolant reuse passage; and a first radially extending turn passage coupling the first rearward passage and the first radially spreading return passage; and a first set of obstructions positioned in the first radially extending turn passage. 
     Two or more aspects described in this disclosure, including those described in this summary section, may be combined to form implementations not specifically described herein. 
     The details of one or more implementations are set forth in the accompanying drawings and the description below. Other features, objects, and advantages will be apparent from the description and drawings, and from the claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       These and other features of this disclosure will be more readily understood from the following detailed description of the various aspects of the disclosure taken in conjunction with the accompanying drawings that depict various embodiments of the disclosure, in which: 
         FIG.  1    shows a schematic cross-sectional view of an illustrative turbomachine in the form of a gas turbine system; 
         FIG.  2    shows a cross-sectional view of an illustrative gas turbine assembly with a three-stage turbine that may be used with the gas turbine system in  FIG.  1   ; 
         FIG.  3    shows a perspective view of an illustrative turbomachine blade in the form of a turbine rotor blade of the type in which embodiments of the disclosure may be employed; 
         FIG.  4    shows a schematic top-down view of cooling circuit(s) in a trailing edge of an airfoil body such as of the turbine rotor blade of  FIG.  3   , according to embodiments of the disclosure; 
         FIG.  5    shows a schematic perspective view of cooling circuits of  FIG.  4    apart from an airfoil body, according to embodiments of the disclosure; 
         FIG.  6    shows a schematic side view of a plurality of first cooling circuits, taken along view line  6 - 6  in  FIG.  4   ; 
         FIG.  7    shows a schematic side view of a plurality of second cooling circuits, taken along view line  7 - 7  in  FIG.  4   ; 
         FIG.  8    shows a schematic side view of a first cooling circuit and a second cooling circuit, as in  FIGS.  6  and  7   , overlaid together; 
         FIG.  9    shows a top-down cross-sectional view of the cooling circuits across view line  9 - 9  in  FIG.  8   ; 
         FIG.  10    shows a top-down cross-sectional view of the cooling circuits across view line  10 - 10  in  FIG.  8   ; 
         FIG.  11    shows a top-down cross-sectional view of the cooling circuits across view line  11 - 11  in  FIG.  8   ; 
         FIG.  12    shows a top-down cross-sectional view of the cooling circuits across view line  12 - 12  in  FIG.  8   ; 
         FIG.  13    shows a top-down cross-sectional view of the cooling circuits across view line  13 - 13  in  FIG.  8   ; 
         FIG.  14    shows a top-down cross-sectional view of the cooling circuits across view line  14 - 14  in  FIG.  8   ; 
         FIG.  15    shows a perspective view of a blade including a cutout for filling with a coupon including cooling circuit(s), according to embodiments of the disclosure; 
         FIG.  16    shows a perspective view of a blade including the coupon including cooling circuit(s), according to embodiments of the disclosure; 
         FIG.  17    shows a perspective view of an illustrative turbomachine blade in the form of a turbine nozzle of the type in which embodiments of the disclosure may be employed; and 
         FIG.  18    shows a schematic cross-sectional view of a conventional trailing edge overlaid with a trailing edge according to embodiments of the disclosure. 
     
    
    
     It is noted that the drawings of the disclosure are not necessarily to scale. The drawings are intended to depict only typical aspects of the disclosure and therefore should not be considered as limiting the scope of the disclosure. In the drawings, like numbering represents like elements among the drawings. 
     DETAILED DESCRIPTION 
     As an initial matter, in order to clearly describe the subject matter of the current disclosure, it will become necessary to select certain terminology when referring to and describing relevant machine components within a turbomachine and/or a turbomachine blade. To the extent possible, common industry terminology will be used and employed in a manner consistent with its accepted meaning. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. Those of ordinary skill in the art will appreciate that often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single part. 
     In addition, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the onset of this section. These terms and their definitions, unless stated otherwise, are as follows. As used herein, “downstream” and “upstream” are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine engine or, for example, the flow of air through the combustor or coolant through one of the turbine&#39;s component systems. The term “downstream” corresponds to the direction of flow of the fluid, and the term “upstream” refers to the direction opposite to the flow (i.e., the direction from which the flow originates). The terms “forward” and “aft,” without any further specificity, refer to directions, with “forward” referring to the front or compressor end of the engine, and “aft” referring to the rearward section of the turbomachine. In context herein, “forward” refers to the leading edge of a turbomachine blade, and “aft” or “rear” refers to the trailing edge of a turbomachine blade. 
     It is often required to describe parts that are disposed at differing radial positions with regard to a center axis. The term “radial” refers to movement or position perpendicular to an axis. For example, if a first component resides closer to the axis than a second component, it will be stated herein that the first component is “radially inward” or “inboard” of the second component. If, on the other hand, the first component resides further from the axis than the second component, it may be stated herein that the first component is “radially outward” or “outboard” of the second component. The term “axial” refers to movement or position parallel to the axis of rotation of the turbine system, or in a chordal direction between leading and trailing edges of an airfoil. Finally, the term “circumferential” refers to movement or position around an axis. It will be appreciated that such terms may be applied in relation to the center axis of the turbine. In the figures (see, e.g., the legend in  FIG.  3   ), an axial orientation is referenced with an “A”; a radial orientation is referenced with an “R”; and a circumferential orientation (about axis A) is referenced with a “C”. 
     In addition, several descriptive terms may be used regularly herein, as described below. The terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises,” “comprising,” “including,” and/or “having,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and/or groups thereof. “Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur or that the subsequently described component or element may or may not be present, and that the description includes instances where the event occurs or the component is present and instances where it does not or is not present. 
     Where an element or layer is referred to as being “on,” “engaged to,” “connected to” or “coupled to” another element or layer, it may be directly on, engaged to, connected to, or coupled to the other element or layer, or intervening elements or layers may be present. In contrast, when an element is referred to as being “directly on,” “directly engaged to,” “directly connected to” or “directly coupled to” another element or layer, there are no intervening elements or layers present. Other words used to describe the relationship between elements should be interpreted in a like fashion (e.g., “between” versus “directly between,” “adjacent” versus “directly adjacent,” etc.). As used herein, the term “and/or” includes any and all combinations of one or more of the associated listed items. 
     As indicated above, the disclosure provides a turbomachine blade and a coupon for a turbomachine blade. The turbomachine blade may include an airfoil body having a pressure side and a suction side connected by a leading edge and a trailing edge, a coolant feed passage defined in the airfoil body, and a coolant reuse (collection) passage defined in the airfoil body. The blade may also include a first cooling circuit defined in the airfoil body. The first cooling circuit may include a rearward passage extending toward the trailing edge from and fluidly coupled to the coolant feed passage, and a radially spreading return passage extending away from the trailing edge toward and fluidly coupled to the coolant reuse passage. The first cooling circuit may also include a radially extending turn passage coupling the rearward passage and the radially spreading return passage. A first set of obstructions may be positioned in the radially extending turn passage. 
     The obstructions, created through additive manufacturing, have a density that allows a lower coolant flow rate and creates sufficient back pressure to allow some of the coolant to exit through vent openings in the trailing edge. The obstructions in the turn passage also provide additional structural strength and allow the trailing edge to have a sharper turn and use thinner walls, thus improving aerodynamic performance of the blade. Coolant not exiting through the vent openings can be reused, for example, for film cooling an exterior surface of the airfoil body or for other purposes. 
     A second cooling circuit may also be provided, e.g., on a pressure side of the airfoil body, to shield parts of the first cooling circuit from a heat load, thus improving the effectiveness of coolant in the first cooling circuit. 
       FIG.  1    shows a schematic illustration of an illustrative turbomachine  100  in the form of a combustion or gas turbine system. Turbomachine  100  includes a compressor  102  and a combustor  104 . Combustor  104  includes a combustion region  105  and a fuel nozzle assembly  106 . Turbomachine  100  also includes a turbine  108  and a common compressor/turbine shaft  110  (sometimes referred to as rotor  110 ). In one embodiment, the combustion turbine system is a 7 HA or 9 HA engine, commercially available from General Electric Company, Greenville, S.C. The present disclosure is not limited to any one particular GT system and may be implemented in connection with other engines including, for example, the other HA, F, B, LM, GT, TM and E-class engine models of General Electric Company and engine models of other companies. Further, the teachings of the disclosure are not necessarily applicable to only a GT system and may be applied to other types of turbomachines, e.g., steam turbines, jet engines, compressors, etc. 
     In operation, air flows through compressor  102 , and compressed air is supplied to combustor  104 . Specifically, the compressed air is supplied to fuel nozzle assembly  106  that is integral to combustor  104 . Assembly  106  is in flow communication with combustion region  105 . Fuel nozzle assembly  106  is also in flow communication with a fuel source (not shown in  FIG.  1   ) and channels fuel and air to combustion region  105 . Combustor  104  ignites and combusts fuel. Combustor  104  is in flow communication with turbine  108  within which gas stream thermal energy is converted to mechanical rotational energy. Turbine  108  is rotatably coupled to and drives rotor  110 . Compressor  102  also is rotatably coupled to rotor  110 . In the illustrative embodiment, there are a plurality of combustors  104  and fuel nozzle assemblies  106 . 
       FIG.  2    shows a cross-sectional view of an illustrative turbine  108  of turbomachine  100  ( FIG.  2   ) with three turbine stages that may be used with the gas turbine system in  FIG.  2   . Each turbine stage of turbine  108  includes a row of stationary blades  112  coupled to a stationary casing of turbomachine  100  and positioned axially adjacent a rotating row of blades  114 . Row of blades  112  includes stationary blades or nozzles  116  (vanes). Each nozzle  116  may be held in turbine assembly  108  by a radially outer platform  118  and a radially inner platform  120 . Row of rotating blades  114  in turbine  108  includes rotating blades  122  coupled to rotor  110  and rotating with the rotor  110 . Rotating blades  122  include a radially inward platform  124  (at root of blade) coupled to rotor  110  and, optionally, may include a radially outward tip shroud  126  (at tip of blade). As used herein, the term “blade” shall refer collectively to stationary blades for vanes or nozzles  116  and rotating blades  122 , unless otherwise stated. 
       FIG.  3    is a perspective view of a blade  130  in the form of a turbine rotor blade  122  of the type in which embodiments of the present disclosure may be employed. Blade  130  includes a root  132  by which blade  130  attaches to rotor  110  ( FIG.  2   ). Root  132  may include a dovetail configured for mounting in a corresponding dovetail slot in the perimeter of the rotor disc. Root  132  may further include a shank that extends between the dovetail and a platform  134 , which is disposed at the junction of an airfoil body  136  and root  132  and defines a portion of the inboard boundary of the flow path through turbine  108 . It will be appreciated that airfoil body  136  is the active component of rotor blade  130  that intercepts the flow of working fluid and induces the rotor disc to rotate. 
     Airfoil body  136  of blade  130  includes a pressure side  140 , i.e., a concave pressure side (PS) outer wall, and a circumferentially or laterally opposite suction side  142 , i.e., a convex suction side (SS) outer wall, extending axially between opposite leading and trailing edges  144 ,  146  respectively. Pressure side  140  and suction side  142  are connected by leading edge  144  and trailing edge  146  and also extend in the radial direction from platform  134  to an outboard tip  148 . Outboard tip  148  is shown without a tip shroud (e.g., tip shroud  126  in  FIG.  2   ). A radially extending coolant feed/reuse circuit  150  may extend between walls  140 ,  142 . While blade  130  of this example is a turbine rotor blade  122  ( FIGS.  2 - 3   ), it will be appreciated that the present disclosure also may be applied to other types of blades within turbine engine  100 , including turbine nozzles  116  ( FIGS.  2  and  17   ) (vanes). The usage of rotor blades in the several embodiments described herein is merely illustrative unless otherwise stated. 
       FIGS.  4 - 14    show various views of cooling circuit(s) defined in airfoil body  136  and, in particular, in trailing edge  146 , according to embodiments of the disclosure.  FIG.  4    shows a schematic top-down view an embodiment of a first cooling circuit  200  and an optional second cooling circuit  202 .  FIGS.  5 - 7    show the portion of cooling circuits  200  and/or  202  in trailing edge  146  within a non-limiting position in blade  130  (designated as position “B” in  FIG.  3   ). More particularly,  FIG.  5    shows a schematic perspective view of cooling circuits  200 ,  202  apart from airfoil body  136  ( FIG.  4   );  FIG.  6    shows a schematic side view of an embodiment of a plurality of first cooling circuits  200  along view line  6 - 6  in  FIG.  4   ; and  FIG.  7    shows a schematic side view of a plurality of second cooling circuits  202  along view line  7 - 7  in  FIG.  4   .  FIG.  8    shows a schematic side view of first cooling circuit  200  and second cooling circuit  202  overlaid together; and  FIGS.  9 - 14    show top-down cross-sectional views across view lines  9 - 9 ,  10 - 10 ,  11 - 11 ,  12 - 12 ,  13 - 13  and  14 - 14  in  FIG.  8   , respectively. 
     Blade  130  ( FIG.  3   ) may include one or more coolant feed passages  204  defined in airfoil body  136  for delivering a coolant to cooling circuit(s)  200 ,  202 . Coolant feed passage(s)  204  may include any, typically radially extending, passage configured to deliver coolant for use in cooling, for example, trailing edge  146 . Coolant can be any now known or later developed coolant used in a turbomachine blade such as, but not limited to, compressed air from compressor  102 , which may be delivered to coolant feed passage(s)  204  through various plenums or casings of the turbomachine and/or blades thereof. Coolant feed passages  204  may be part of a coolant feed/reuse circuit  150  ( FIG.  3   ). 
     Blade  130  ( FIG.  3   ) may also include one or more a coolant reuse passages  206  defined in airfoil body  136  for collecting coolant from cooling circuit(s)  200 ,  202  for reuse in cooling other parts of the blade. Coolant reuse passage(s)  206  may include any typically radially extending passage configured to collect coolant from cooling circuit(s)  200 ,  202 . Coolant reuse passage(s)  206  may route used coolant to other parts of airfoil body  136  or parts of blade  130 , e.g., tip, tip shroud, etc., or it may route used coolant to exterior surfaces of airfoil body  136 , e.g., for film cooling. In the latter case, as shown in  FIG.  4   , coolant reuse passage(s)  206  may include vent openings  208  to pressure side  140  and/or suction side  142 . 
     In some of the drawings, two coolant feed passages  204  are shown, one for first cooling circuit  200  and one for second cooling circuit  202 . Other drawings, such as  FIG.  8   , show a shared coolant feed passage  204  and a shared coolant reuse passage  206 . It is emphasized that any number of coolant feed or reuse passages  204 ,  206  can be employed. 
     First cooling circuit  200  is defined in airfoil body  136  or, as will be described, in a coupon body  270  ( FIG.  16   ). In the example shown, first cooling circuit  200  is adjacent suction side  142  of airfoil body  136 , which is cooler than pressure side  140  of airfoil body  136  during operation in a turbomachine. First cooling circuit  200  may include a first rearward (feed or inlet) passage  220  extending toward trailing edge  146  from and fluidly coupled to coolant feed passage  204 . First rearward passage  220 , which extends generally axially, can have any tubular cross-sectional shape, e.g., circular. First cooling circuit  200  also includes a first radially spreading return passage  222  extending away from trailing edge  146  toward and fluidly coupled to first coolant reuse passage  206 . As shown best in  FIGS.  5 ,  6  and  8   , radially spreading return passage  222  has a radial extent R 1  that is significantly greater than a radial extent R 2  of rearward passage  220 , e.g., greater than 3 times. 
     As shown in the lower return passage in  FIG.  6   , a set of obstructions  224 , such as a pin or fin bank, may be optionally positioned in first radially spreading return passage  222  to increase cooling, improve structural integrity, and/or control back pressure. 
     As shown best in  FIGS.  4  and  5   , return passage  222  may be partially or completely between rearward passage  220  and a hot gas path (HGP)  226  about suction side  142  of airfoil body  136 , which limits the amount of energy picked up by coolant before entering trailing edge  146 . In this manner, most of the coolant&#39;s energy is used in trailing edge  146  rather than prior to trailing edge  146 . Coolant returning in return passage  222  may be reused in any manner, e.g., film cooling for suction side  142  and/or pressure side  140 . As shown in  FIGS.  5 ,  6  and  8   , first rearward passage  220  is radially offset from first radially spreading return passage  222  along a radial axis R of airfoil body  136 . 
     As shown in  FIGS.  4 ,  5 ,  6 ,  8 - 14   , first cooling circuit  200  also may include a first radially extending turn passage  230  (hereafter “turn passage  230 ”) coupling first rearward passage  220  and first radially spreading return passage  222 . Turn passage  230  may extend any radial extent R 3  ( FIG.  8   ) necessary to fluidly couple rearward passage  220  and return passage  222 . As shown for example in  FIGS.  5  and  9 - 12   , turn passage  230  may have a first circumferential width W 1  at a forward end thereof that is greater than a second circumferential width W 2  at an aft end thereof. The elliptical shape near the aft end (closest to trailing edge  146 ) has a length-to-width (L/W 2 )(see e.g.,  FIG.  9   ) ratio of between 1:1 and 4:1, inclusive of end values. In this manner, coolant passes as close as possible to trailing edge  146  in turn passage  230 , and a shape of trailing edge  146  can be more narrow and/or pointed compared to previous blades to improve aero-performance. 
     As illustrated for example in  FIG.  8   , return passage  222  may fluidly couple to turn passage  230  via a coupling passage  232  that is radially smaller (R 4 ) than return passage  222  (R 2 ), i.e., it has a smaller cross-sectional area. A coolant flow (dark arrows) may travel axially rearwardly through rearward passage  220  from feed passage  204 , radially outward in turn passage  230  and then axially forward starting in coupling passage  232  and may then expand radially as it flows axially forward in return passage  220  to reuse passage  206 . Coolant flows radially outward in turn passage  230  and over a separating wall  236  between turn passage  230  and return passage  222 . 
     Referring to  FIGS.  6  and  8   , a first set of obstructions  238  may be positioned in turn passage  230 . Obstructions  238  can take the form of any structure typically used to improve structural integrity of a passage, create back pressure and/or improve heat transfer of a coolant flowing therethrough. In certain embodiments, as shown in the top turn passage  230  in  FIG.  6   , obstructions  238  may be cylindrical pins having a circular cross-section. In other embodiments, as shown in the bottom turn passage  230  in  FIG.  6    and in  FIG.  8   , obstructions  238  may be polygonal pegs having polygonal cross-sections, e.g., rectangular, square, pentagonal, etc. The size and spacing of obstructions  238  that creates a specific density of obstructions can be selected to control, for example, heat transfer and back pressure in turn passage  230 . 
     In particular, obstruction density may be increased to increase back pressure, which allows vent passages  240  (described herein) through trailing edge  146  to provide more direct exit of coolant, increases total flow rate though the cooling features, and increases cold side surface area for heat transfer. Hence, obstructions  238  enhance heat transfer and increase the surface area available to transfer energy to the coolant. Obstructions  238  also act as a metering area, allowing, for example, an increased number of vent passages  240  to be used on pressure side  140 , increasing coolant film coverage. Obstructions  238  also add to the structural integrity of trailing edge  146 . 
     In one non-limiting example, obstructions  238  were square and had side dimensions of 0.305-1.524 millimeters (0.012 to 0.060 inches) with spacing ranging from 1.07-1.73 times the side dimensions in a transverse-to-flow direction and 0.41-1.45 times the side dimensions in a flow direction (see arrows). In another non-limiting example, obstructions  238  had circular cross-sectional diameters of 0.305-1.067 (0.012-0.042 inches) with spacings ranging from 1.2-3 times the diameter in the transverse-to-flow direction and 1.1-1.7 times the diameter in the flow direction. In any event, the density of set of obstructions  238  can be selected to control, for example, structural strength and/or back pressure. The number, shapes and/or sizes of obstructions  238  in turn passages  230  (and other obstructions, described herein) may be the same throughout a given blade  130 , or they may vary depending on, for example, radial location, turn passage size or shape, required heat transfer, required structural strength, number of vent passages  240  to be used, among other factors. 
     As shown in  FIGS.  6 ,  8 - 11 ,  13  and  14   , blade  130  may also include a plurality of vent passages  240  extending from radially extending turn passage  230  through trailing edge  146  of airfoil body  136 . Coolant can vent through trailing edge  146  by way of vent passages  240 . Any number of vent passages  240  may extend from turn passage  230  and may be at any angle/orientation, cross-sectional size or shape, and number, desired to create a particular coolant flow and/or to eliminate the need for any minimum trailing edge thickness based on manufacturing tolerances. Vent passages  240  may be angled toward pressure side  140  and/or suction side  142 . Vent passages  240  provide a more direct exit for coolant flow, increasing the total flow rate through trailing edge  146  and also increasing the cold-side surface area for heat transfer. 
     Referring to  FIGS.  4 ,  5 ,  7 ,  8  and  11 - 14   , blade  130  may also optionally include second cooling circuit  202  defined in airfoil body  136 . Second cooling circuit  202  is defined in airfoil body  136  or, as will be described, in coupon body  270  ( FIG.  16   ). In the example shown, second cooling circuit  202  is adjacent pressure side  140  of airfoil body  136 . Hence, second cooling circuit  202  provides cooling near pressure side  140  of airfoil body  136 , which is hotter than suction side  142  of airfoil body  136  during operation of a turbomachine. As shown in  FIGS.  4  and  5   , first cooling circuit  200  is circumferentially offset from the second cooling circuit  202  in airfoil body  136 . Hence, second cooling circuit  202  also acts as a buffer to protect coolant in first cooling circuit  200  from excessive heat transfer (e.g., from pressure side  140  of airfoil body  136 ) prior to reaching trailing edge  146 . That is, second cooling circuit  202  shields incoming coolant in first cooling circuit  200  from heat transfer in addition to cooling pressure side  140  of airfoil body  136 . 
     Second cooling circuit  202  is somewhat similar in shape to first cooling circuit  200 . Second cooling circuit  202  may include a second rearward (feed or inlet) passage  250  extending toward trailing edge  146  (but not reaching it) from and fluidly coupled to coolant feed  204 . Second rearward passage  250 , which extends generally axially, can have any tubular cross-sectional shape, e.g., circular. Coolant feed  204  coupled to second rearward passage  250  may be a separate coolant feed (see e.g.,  FIGS.  4 - 5   ) from that of first cooling circuit  200 , or it may be a shared coolant feed  204  (see e.g.,  FIG.  8   ). Second cooling circuit  202  also includes a second radially spreading return passage  252  extending away from trailing edge  146  toward and fluidly coupled to a second coolant reuse passage  206  defined in airfoil body  136 . Second coolant reuse passage  206  may be a separate coolant reuse passage  206  (see e.g.,  FIGS.  4 - 5   ) from that of first cooling circuit  200 , or it may be a shared coolant reuse passage  206  (see e.g.,  FIG.  8   ). 
     As shown best in  FIGS.  5 ,  7  and  8   , radially spreading return passage  252  has a radial extent R 5  that is significantly greater than a radial extent R 6  of rearward passage  250 , e.g., greater than 3 times. As shown in the lower return passage in  FIG.  7   , a set of obstructions  256 , such as a pin or fin bank, may be positioned in second radially spreading return passage  252  to increase cooling, improve structural integrity, and/or control back pressure. As shown best in  FIGS.  4  and  5   , return passage  252  may be partially or completely between rearward passage  250  and a hot gas path (HGP)  258  about pressure side  140  of airfoil body  136 , which limits the amount of energy picked up by coolant in second cooling circuit  202  before entering trailing edge  146 . In this manner, most of the coolant&#39;s energy is used in trailing edge  146  rather than prior to trailing edge  146 . Coolant returning in return passage  252  may be reused in any manner, e.g., film cooling for suction side  142  and/or pressure side  140 . As shown in  FIGS.  5 ,  6  and  8   , second rearward passage  250  is radially offset from second radially spreading return passage  252  along a radial axis R of airfoil body  136 . 
     Second coolant circuit  202  also includes a second radially extending turn passage  260  (hereafter “turn passage  260 ”) coupling second rearward passage  250  and second radially spreading return passage  252 . Turn passage  260  may extend any radial extent R 7  ( FIGS.  7 ,  8   ) necessary to fluidly couple rearward passage  250  and return passage  252  of second cooling circuit  202 . As shown best in  FIGS.  5  and  8   , second rearward passage  250  is radially offset from second radially spreading return passage  252  along radial axis R of airfoil body  136 . As illustrated for example in  FIG.  8   , return passage  252  may fluidly couple to turn passage  250  via a coupling passage  262  that is radially smaller (R 8 ) than return passage  252  (R 5 ), i.e., it has a smaller cross-sectional area. A coolant flow (dark arrows) may travel axially rearwardly through rearward passage  250  from feed passage  204 , radially outward in turn passage  260  and then axially forward starting in coupling passage  262  before expanding as it flows axially forward in return passage  252  to reuse passage  206 . Coolant flows radially outward in turn passage  260  and over a separating wall  264  between turn passage  260  and return passage  252 . 
     With reference to  FIGS.  6  and  7   , any number of first or second cooling circuits  200 ,  202  may be radially positioned in blade  130 . That is, blade  130  may include a plurality of the first cooling circuits  200  radially spaced in airfoil body  136  and a plurality of second cooling circuits  202  radially spaced in airfoil body  136 . While two of each are shown in  FIGS.  6  and  7   , any number may be used, i.e., one of each or three or more. 
     As noted, embodiments of the disclosure can be used in a turbomachine blade  130  or in a coupon  270  ( FIG.  16   ) that replaces a part of a turbomachine blade, i.e., a part of an airfoil thereof. With reference to  FIG.  15   , embodiments of a coupon  270  (in dashed lines) for use with a preexisting turbomachine blade  272 , such as a rotor blade  130  ( FIG.  3   ), will now be described.  FIG.  15    shows a perspective view of a preexisting turbine rotor blade  272  (hereinafter “blade  272 ”). Blade  272  may include external and internal structure as described relative to turbine rotor blade  130  of  FIG.  3   . Blade  272  includes airfoil body  136 . Airfoil body  136  of blade  272  includes pressure side  140 , i.e., a concave pressure side (PS) outer wall, and circumferentially or laterally opposite suction side  142 , i.e., a convex suction side (SS) outer wall, extending axially between opposite leading and trailing edges  144 ,  146 , respectively. Pressure side  140  and suction side  142  are connected by leading edge  144  and trailing edge  146  and extend in the radial direction from platform  134  to an outboard tip  148 . Outboard tip  148  is shown without a tip shroud (e.g., tip shroud  126  in  FIG.  2   ). A radially extending coolant feed/reuse circuit  150  may extend between walls  140 ,  142 . 
     Blade  272  is shown with a cutout  280  positioned along the aft end of the airfoil body  136  (that is, a portion encompassing trailing edge  146 ). As illustrated in the example in  FIG.  15   , cutout  280  is a predetermined area (shown by dashed lines) that encompasses trailing edge  146  and that is removed from airfoil body  136 , i.e., a predetermined portion extending forward from trailing edge  146 . Cutout  280  can be removed by any now known or later developed metal cutting technique, e.g., welding torch, electrical discharge machining (EDM), laser cutting, water jet cutting, etc. As illustrated, cutout  280  includes most, if not all, of a radial extent of airfoil body  136 . It is emphasized, however, that cutout  280  can include any portion of airfoil body  136  within which first and/or second cooling circuits  200 ,  202  ( FIGS.  4 - 14   ) in trailing edge  146  may be desired. 
     As shown in  FIG.  16   , coupon  270  is coupled in cutout  280  ( FIG.  15   ) to replace a predetermined area ( FIG.  15   ) of airfoil body  136  that encompasses trailing edge  146 . In accordance with embodiments of the disclosure, coupon  270  includes the structure described herein relative to first cooling circuits  200  and, where desired, second cooling circuits  202 . More particularly, coupon  270  includes a coupon body  284  including first cooling circuit(s)  200  ( FIGS.  4 - 6  and  8 - 14   ) defined therein and, optionally, second cooling circuit(s)  202  ( FIGS.  4 - 5 ,  7  and  11 - 14   ). First cooling circuit  200  and second cooling circuit  202  may be arranged as described herein. Coupon  270  can be the same size as cutout  280  or larger or smaller. The replacement of at least a portion of trailing edge  146  with coupon  270  can enhance performance of existing blades  272  by reducing coolant flow out of trailing edge  146 . Coolant used by first and second cooling circuits  200 ,  202  can be obtained from any coolant feed  204  (see e.g.,  FIGS.  4 - 5 ,  8   ) and may be reused as described herein. 
       FIG.  17    is a perspective view of a blade  130  in the form of a turbine nozzle  116  of the type in which embodiments of the present disclosure may be employed. Nozzle  116  may be held in turbine assembly  108  ( FIG.  2   ) by radially outer platform  118  and radially inner platform  120 . It will be appreciated that airfoil body  136  is the active component of blade  130  (nozzle  116 ) that intercepts the flow of working fluid and directs the flow where desired. Airfoil body  136  of nozzle  116  includes pressure side  140 , i.e., a concave pressure side (PS) outer wall, and a circumferentially or laterally opposite suction side  142 , i.e., a convex suction side (SS) outer wall, extending axially between opposite leading and trailing edges  144 ,  146 , respectively. Pressure side  140  and suction side  142  are connected by leading edge  144  and trailing edge  146  and extend in the radial direction from radially inner platform  120  to radially outer platform  118 . A radially extending coolant feed/reuse circuit  150  may extend between walls  140 ,  142 . Nozzle  116  may include first cooling circuit  200  ( FIGS.  4 - 6  and  8 - 14   ), and optionally second cooling circuit  202  ( FIGS.  4 - 5 ,  7  and  11 - 14   ). A coupon  270 , as described relative to  FIGS.  15  and  16   , may be applied to nozzle  116  in a similar fashion as described relative to turbine rotor blade  122 . 
     Blade  130  (rotating blades and stationary vanes) or coupon  270  may include any metal or metal compound capable of withstanding the environment in which used. Blade  130  and coupon  270  may be made using any now known or later developed manufacturing technique. However, additive manufacturing allows for blade  130  and coupon  270  to be formed with greatly minimized sizes and shapes, e.g., smaller obstructions and thinner walls of airfoil body  136 , many of which improve aerodynamic performance. As used herein, additive manufacturing (AM) may include any process of producing an object through the successive layering of material rather than the removal of material, which is the case with conventional processes. Additive manufacturing can create complex geometries without the use of any sort of tools, molds or fixtures, and with little or no waste material. Instead of machining components from solid billets of metal, much of which is cut away and discarded, the material used in additive manufacturing is only what is required to shape the part. Additive manufacturing processes may include, but are not limited to: 3D printing, rapid prototyping (RP), direct digital manufacturing (DDM), binder jetting, selective laser melting (SLM) and direct metal laser melting (DMLM). In the current setting, DMLM has been found advantageous. 
     Embodiments of the disclosure can improve aerodynamic efficiency of turbomachine blades by providing a trailing edge having a sharper turn than conventional blades.  FIG.  18    shows a schematic cross-sectional view of a conventional trailing edge  300  overlaid with a trailing edge  302 , according to embodiments of the disclosure. As shown, trailing edge  302  is more pointed or sharper than conventional trailing edge  300 , the latter of which has a more uniform radius R 9  and more square profile. Hence, trailing edge  302  is more highly elliptical. More particularly, a conventional trailing edge  300  typically has an ellipse ratio of less than or equal to 1 as illustrated by the ellipse  304 —see major and minor axes labeled  304 maj,  304 min, respectively. (Ellipse ratio is equal to major axis divided by minor axis. The major axis as defined herein is generally along the mean camber line  305  of airfoil body  136 , it is not perpendicular to the mean camber line  305 .) The ellipse ratio may be the result of manufacturing of the trailing edge and/or the provision of thermal barrier coatings thereon. 
     Currently, an ellipse ratio of greater than 1 is challenging to manufacture because it is difficult to sufficiently cool. However, in certain embodiments of the disclosure, an ellipse ratio of trailing edge  302  can be between 1.1 to 4—see major and minor axes labeled  302 maj,  302 min, respectively. In other embodiments, the ellipse ratio of trailing edge  302  can be between 1.1 to 3. In further embodiments, the ellipse ratio of trailing edge  302  can be between 1.1 to 2. In yet other embodiments, the ellipse ratio of trailing edge  302  can be between 1.1-1.5. For purposes of evaluation, a location of trailing edge  302  may be defined based on where airfoil body  136  transitions from the more linear pressure side  140  or suction side  142  to have more curvature, i.e., with a large gradient in curvature typical of a trailing edge compared to the rest of airfoil body  136 . The transition in curvature may be identified, for example, using a curvilinear combs graphical analysis tool available in any now known or later developed computer aided graphics (CAD) design system. In  FIG.  18   , illustrative transition points for conventional trailing edge  300  are labeled  306 , and those for trailing edge  302  are labeled  308 . 
     Embodiments of the disclosure can also improve aerodynamic efficiency of turbomachine blades, e.g., by using thinner walls, with reduced coolant flow to reduce trailing edge temperatures. In addition, the obstructions in the turn passage also provide additional structural strength. Where a coupon is used to provide the cooling circuits to a preexisting blade, the coupons can provide internal cooling structures not previously present in the blade, thus providing improved cooling and aero-performance and lengthening a lifespan of the part. 
     Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about,” “approximately” and “substantially,” are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. “Approximately,” as applied to a particular value of a range, applies to both end values and, unless otherwise dependent on the precision of the instrument measuring the value, may indicate +/−10% of the stated value(s). 
     The corresponding structures, materials, acts, and equivalents of all means or step plus function elements in the claims below are intended to include any structure, material, or act for performing the function in combination with other claimed elements as specifically claimed. The description of the present disclosure has been presented for purposes of illustration and description but is not intended to be exhaustive or limited to the disclosure in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the disclosure. The embodiment was chosen and described in order to best explain the principles of the disclosure and the practical application and to enable others of ordinary skill in the art to understand the disclosure for various embodiments with various modifications as are suited to the particular use contemplated.