BACKGROUND 1. Field

[0001] This invention relates generally to an airfoil in a turbine engine, and in particular, to a trailing edge cooling feature incorporated in a turbine airfoil.

2. Description of the Related Art

[0002] In gas turbine engines, compressed air discharged from a compressor section and fuel introduced from a source of fuel are mixed together and burned in a combustion section, creating combustion products defining a high temperature working gas. The working gas is directed through a hot gas path in a turbine section of the engine, where the working gas expands to provide rotation of a turbine rotor. The turbine rotor may be linked to an electric generator, wherein the rotation of the turbine rotor can be used to produce electricity in the generator.

[0003] In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as airfoils, e.g., stationary vanes and rotating blades within the turbine section, must be cooled with cooling fluid, such as air discharged from a compressor in the compressor section, to prevent overheating of the components.

[0004] Effective cooling of turbine airfoils requires delivering the relatively cool air to critical regions such as along the trailing edge of a turbine blade or a stationary vane. The associated cooling apertures may, for example, extend between an upstream, relatively high pressure cavity within the airfoil and one of the exterior surfaces of the turbine blade. Blade cavities typically extend in a radial direction with respect to the rotor and stator of the machine.

[0005] Airfoils commonly include internal cooling channels which remove heat from the pressure sidewall and the suction sidewall in order to minimize thermal stresses. Achieving a high cooling efficiency based on the rate of heat transfer is a significant design consideration in order to minimize the volume of coolant air diverted from the compressor for cooling. However, the relatively narrow trailing edge portion of a gas turbine airfoil may include, for example, up to about one third of the total airfoil external surface area. The trailing edge is made relatively thin for aerodynamic efficiency. Consequently, with the trailing edge receiving heat input on two opposing wall surfaces which are relatively close to each other, a relatively high coolant flow rate is entailed to provide the requisite rate of heat transfer for maintaining mechanical integrity.

SUMMARY

[0006] An object of the invention is to provide an improved trailing edge cooling feature for a turbine airfoil.

[0007] According to a first aspect of the invention, an airfoil for a turbine engine is provided. The airfoil incorporates a trailing edge cooling feature including a plurality of profiled partition walls connecting the pressure side and the suction side of an outer wall of the airfoil and generally extending from the radially inner end to the radially outer end of the outer wall. The partition walls are spaced apart from each other in the chordal direction to define cooling passages therebetween. Each cooling passage extends from the radially inner end to the radially outer end along a flow path featuring a series of flow turns defined by the profile of adjoining partition walls. The partition walls have coolant impingement openings that fluidically interconnect the cooling passages with each other. The coolant impingement openings on adjacent partition walls are staggered along a radial direction of the turbine engine.

[0008] In at least one embodiment, the outer wall of the airfoil delimits an airfoil interior and includes including a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end and a radially outer end, wherein a chordal direction is defined between the leading edge and the trailing edge. A coolant fluid cavity is located in the airfoil interior and extending generally radially between the radially inner end and the radially outer end. The coolant fluid cavity receives a coolant fluid for cooling the outer wall. The inventive trailing edge cooling feature located downstream of the coolant fluid cavity in the chordal direction.

[0009] The inventive trailing edge cooling feature provides a coolant flow path that is much longer inside the airfoil than in conventional designs, for absorbing more heat from the outer wall of the airfoil, thereby providing more cooling to the outer wall before exiting the airfoil at the trailing edge into the hot gas path. In addition, the inventive design provides serial impingement along the chordal direction which forms a high pressure drop and effective metering of the coolant flow rate through the trailing edge cooling system. Furthermore, as a result of the radial staggering of the coolant impingement openings, the coolant, after flowing through each impingement opening, flows both radially outwardly and radially inwardly along the profiled flow path in the cooling passage before entering into the next serial impingement openings. The inventive design thus provides a combination of impingement and convection cooling in a serial arrangement before exiting the airfoil trailing edge [0010] To take advantage of the increased cooling provided by the long flow path of the coolant, the number of impingement openings in each partition wall may be reduced, which would reduce the amount of coolant flow. The inventive trailing edge cooling feature is therefore particularly suitable for low cooling flow design airfoils, for example, first row vanes and blades. [0011] In one embodiment, the coolant impingement openings of adjacent first and second partition walls are staggered in the radial direction in a manner that the flow path of a cooling passage defined between the adjacent first and second partition walls exhibits at least one flow turn between a coolant impingement opening on the first partition wall and a coolant impingement opening on the second partition wall which is radially most proximate to the coolant impingement opening on the first partition wall. The at least one flow turn between serial impingement openings provides a longer flow path for the coolant fluid post impingement through the first partition wall to provide more heat transfer before entering the next impingement opening at the second partition wall. [0012] In one embodiment, to provide increased turbulence in the cooling passages, the coolant impingement openings open into flow turns of the cooling passages.

[0013] In certain embodiments, at least one of the cooling passages comprises a flow path having periodic profile along the radial direction, defined by alternating peaks and valleys spaced apart in the radial direction, wherein the peaks and valleys form the flow turns. Such a periodic profile with alternating peaks and valleys provides a particularly long flow path within the cooling passages to provide a significantly increased heat transfer surface with the partition walls. An example of a periodic profile is a wavy profile. Another example of a periodic profile may be a triangular profile.

[0014] In a preferred embodiment, the flow turns in the cooling passages comprise smooth curves. An advantage of this feature is that it avoids or at least minimizes flow separation of the coolant at the turns, thereby increasing the rate of convective heat transfer. Another advantage of this feature is that it provides increased mechanical strength to the airfoil body to withstand centrifugal loads during operation of the turbine engine. Yet another advantage of the feature lies in the fact it provides easier casting by reducing stress concentration that would otherwise be caused by sharp angled turns.

[0015] In one embodiment, the coolant impingement openings are located substantially equidistant from the suction side and the pressure side. In an alternate embodiment, the coolant impingement openings 48A-E are arranged at varying positions in the direction from the pressure side 24 to the suction side 26.

[0016] In one embodiment, to provide improved metering of the coolant flow rate, the coolant impingement openings have different sizes such that a coolant impingement opening on a first partition wall has a larger flow cross-section than that of a coolant impingement opening on a second partition wall at a downstream location from the first partition wall along the chordal direction

[0017] In a further embodiment, the airfoil further comprises a plurality of turbulating features located within at least one of the cooling passages, the turbulating features being provided on the suction side and/or pressure side of the outer wall. The turbulating features increase the heat transfer surface while also providing increased turbulence within the cooling passage.

[0018] In a further embodiment, the airfoil further comprises a plurality of outlet passages located in the outer wall at the trailing edge, the outlet passages receiving coolant fluid from the radially extending cooling passages via one or more of the coolant impingement openings, and discharging the coolant fluid from the trailing edge of the airfoil.

[0019] In a still further embodiment, the airfoil further comprises a coolant fluid channel located between the cooling passages and the outlet passages and extending generally radially between the radially inner end and the radially outer end of the outer wall, the coolant fluid channel receiving coolant fluid from the radially extending cooling passages via one or more of the coolant impingement openings and delivering the coolant fluid to the outlet passages. [0020] In one embodiment, the partition walls are cast integrally with the outer wall using a ceramic casting core.

[0021] According a second aspect of the invention, a casting core is provided. The casting core may be designed for forming a turbine airfoil with the inventive trailing edge cooling as described above.

BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The invention is shown in more detail by help of figures. The figures show preferred configurations and do not limit the scope of the invention. [0023] FIG 1 is a side cross-sectional view of an airfoil assembly wherein a portion of a suction side of the airfoil assembly has been removed to expose a trailing edge cooling feature according to one embodiment of the invention,

[0024] FIG 2 is cross-sectional view of the airfoil assembly of FIG 1 taken along the section line II-II in FIG. 1 ;

[0025] FIG 3 is an enlarged side cross-sectional view of a portion of the airfoil assembly of FIG 1 that includes the trailing edge cooling feature according the illustrated embodiment;

[0026] FIG 4 is a cross-sectional view of the airfoil assembly taken along the section line IV-IV in FIG 3,

[0027] FIG 5 is a cross-sectional view of the airfoil assembly taken along the section line V-V in FIG 4, and

[0028] FIG 6 is an enlarged side cross-sectional view of a portion of a casting core for forming a turbine airfoil with a trailing edge cooling feature according to one embodiment of the invention.

DETAILED DESCRIPTION

[0029] In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

[0030] Referring now to FIG 1, an airfoil assembly 10 constructed in accordance with a first embodiment of the present invention is illustrated. In the embodiment illustrated in FIG 1, the airfoil assembly 10 is a blade assembly comprising an airfoil 12, i.e., a rotatable blade, although it is understood that the cooling concepts disclosed herein could be used in combination with a stationary vane. The airfoil assembly 10 is for use in a turbine section 14 of a gas turbine engine.

[0031] As will be apparent to those skilled in the art, the gas turbine engine includes a compressor section (not shown), a combustor section (not shown), and the turbine section 14. The compressor section includes a compressor that compresses ambient air, at least a portion of which is conveyed to the combustor section. The combustor section includes one or more combustors that combine the compressed air from the compressor section with a fuel and ignite the mixture creating combustion products defining a high temperature working gas. The high temperature working gas travels to the turbine section 14 where the working gas passes through one or more turbine stages, each turbine stage comprising a row of stationary vanes and a row of rotating blades. It is contemplated that the airfoil assembly 10 illustrated in FIG 1 may be included in a first row of rotating blade assemblies or vane assemblies in the turbine section 14, although the incorporation of the same features in other rows of blades and/or vanes may be readily conceived.

[0032] The vane and blade assemblies in the turbine section 14 are exposed to the high temperature working gas as the working gas passes through the turbine section 14. Cooling air from the compressor section may be provided to cool the vane and blade assemblies, as will be described herein.

[0033] As shown in FIG 1, the airfoil assembly 10 comprises the airfoil 12 and a platform assembly 16 that is coupled to a turbine rotor (not shown) and to which the airfoil 12 is affixed. The airfoil 12 comprises an outer wall 18 (see FIGS 1 and 2) that extends span-wise along a radial direction R with respect to a rotation axis A of the turbine engine, and comprises a radially inner end 18A and a radially outer end 18B. The radially inner end 18A is affixed to the platform assembly 16. The radially outer end 18B forms the tip of the airfoil 12.

[0034] Referring to FIG 2, the outer wall 18 further includes a leading edge 20, a trailing edge 22, a concave-shaped pressure side 24 and a convex-shaped suction side 26. A chordal direction C is defined as a direction from the leading edge 20 to the trailing edge 22. It is noted that a portion of the suction side 26 of the airfoil 12 illustrated in FIG 1 has been removed to show selected internal structures within the airfoil 12, as will be described herein.

[0035] As shown in FIG 2, an inner surface 18C of the outer wall 18 defines a hollow interior portion 28, which is delimited by the outer wall 18. The hollow interior portion 28 extends between the pressure and suction sides 24, 26 from the leading edge 20 to the trailing edge 22 and from the radially inner end 18A to the radially outer end 18B. A plurality of rigid spanning structures 30 extend within the hollow interior portion 28 from the pressure side 24 to the suction side 26 and from the radially inner end 18A to the radially outer end 18B to provide structural rigidity for the airfoil 12 and to divide the hollow interior portion 28 into a plurality of sections, which will be described below. The spanning structures 30 may be formed integrally with the outer wall 18. A conventional thermal barrier coating (not shown) may be provided on an outer surface 18D of the outer wall 18 to increase the heat resistance of the airfoil 12, as will be apparent to those skilled in the art.

[0036] In accordance with the present invention, the airfoil assembly 10 is provided with a trailing edge cooling feature 40 for effecting cooling of the airfoil 12 toward the trailing edge 22 of the outer wall 18. As noted above, while the description of the trailing edge cooling feature 40 pertains to a blade assembly, it is contemplated that the concepts of the trailing edge cooling feature 40 of the present invention could be incorporated into a vane assembly.

[0037] As shown in FIGS 1 and 2, the trailing edge cooling feature 40 is located in the hollow interior portion 28 of the outer wall 18 toward the trailing edge 22. A coolant fluid cavity 42 is defined in the outer wall 18 between the pressure side 24 and the suction side 26 and extending generally radially between the radially inner end 18 A and the radially outer end 18B of the outer wall 18. The coolant fluid cavity 42 receives coolant fluid from the platform assembly 16 for cooling the outer wall 18 near the trailing edge 22. The trailing edge cooling feature 40 is positioned downstream of the coolant fluid cavity along the chordal direction C. [0038] As seen in FIGS 1 and 2, the trailing edge cooling feature 40 includes a plurality of cooling passages 46, defined in the interspaces between a plurality of partition walls 44. The partition walls 44 extend from the pressure side 24 to the suction side 26 of the outer wall 18 as shown in FIG 2, and extend span- wise generally in the radial direction R from the radially inner end 18A to the radially outer end 18B along a profiled contour as shown in FIG 1. The partition walls 44 are provided with coolant impingement openings 48 that interconnect the cooling passages 46.

[0039] The trailing edge cooling feature 40 is illustrated in greater detail in the enlarged views shown in FIG 3 and FIG 4. As shown therein, a plurality of partition walls 44A-E are arranged downstream of the coolant fluid cavity 42 in the chordal direction C. By way of example only, the illustrated embodiment shows five partition walls, each being individually designated as 44A, 44B, 44C, 44D, 44E. As shown in FIG 4, each of partition walls 44A-E connects the pressure side 24 with the suction side 26 of the outer wall 18. Further, as shown in FIGS 1 and 3, each of the partition walls 44A-E is profiled, extending generally in the radial direction R from the radially inner end 18A to the radially outer end 18B.

[0040] The term "profiled", in the context of this discussion, implies having a contour that is not a straight line along the radial direction R, but instead having a plurality of angular turns, which may be smooth or sharp, as it extends generally in the radial direction R from the radially inner end 18A to the radially outer end 18B. The angular turns divide the contour into a plurality of sections, which may be straight or curved.

[0041] The profiled partition walls 44A-E are spaced apart from each other in the chordal direction C to define cooling passages 46A, 46B, 46,C, 46D therebetween. Each cooling passage 46A-D therefore has a profile defined by that of a pair of adjoining profiled partition walls 44A-E. For example, the profile of the cooling passage 46A is defined by the profile of an upstream partition wall 44A and that of a downstream partition wall 44B. Each cooling passage 46A-D may therefore be understood to be extending from the radially inner end 18A to the radially outer end 18B (see FIG 1) along a flow path F comprising a series of flow turns T l, T2 defined by the profile of adjoining partition walls 44A-E. [0042] Each of the partition walls 44A-E comprises one or more coolant impingement openings 48A-E that fluidically interconnect the cooling passages 46A- D with each other and with the coolant fluid cavity 42. In the illustrated embodiment, each partition wall 44A-E comprises a plurality of coolant impingement openings 48A-E which are spaced apart from each other in the radial direction R. Each of the coolant impingement openings 48A-E may comprise a through-hole, defining a passage through the respective partition wall 44A-E. As shown in FIGS 3 and 4, the passages defined by the coolant impingement openings 48A-E have at least a component in the chordal direction C. In the illustrated embodiment, the passages defined by the coolant impingement openings 48A-E are parallel to the chordal direction C and have no component in the radial direction R. In an alternate embodiment, one or more of the coolant impingement openings 48A-E may have passages that are angled, i.e., have a component in the chordal direction C as well as a component in the radial direction R. [0043] As shown in FIG 3, the coolant impingement openings 48A, 48B on adjacent partition walls 44A, 44B respectively are staggered along a radial direction R of the turbine engine, i.e. staggered along a span-wise direction of the airfoil 12. A staggering along the radial direction R implies that no two coolant impingement openings 48A-E on adjacent partition walls 44A-E are positioned at the same level along the radial direction R. For example, each of the coolant impingement openings 48A on the partition wall 44A are at different radii in relation to each of the coolant impingement openings 48B on the adjacent partition wall 44B.

[0044] In at least one embodiment, as illustrated in FIGS 3 and 4, the radial staggering between coolant impingement openings 48A-E on adjacent partition walls 44A-E is such as to include at least one flow turn Tl, T2 between the nearest coolant impingement openings 48A-E on adjacent partition walls 44A-E. That is to say that the coolant impingement openings, for example, 48A, 48B of adjacent partition walls 44A, 44B respectively are staggered in the radial direction R in a manner that the flow path F of the cooling passage 46A defined between the adjacent partition walls 44A, 44B exhibits at least one flow turn Tl, T2 between a coolant impingement opening 48A on the partition wall 44A and the nearest coolant impingement opening 48B on the partition wall 44B. This ensures that the coolant flowing through the coolant impingement opening 48A is turned at least once in the cooling passage 46A before exiting the cooling passage 46A via the nearest coolant impingement 48B, thereby providing longer overall flow path for the coolant in the cooling passage 46A.

[0045] In one embodiment, the flow path F of one or more cooling passages 46A- D may have a periodic profile along the radial direction R defined by alternating peaks Tl and valleys T2 spaced apart in the radial direction R. In this case, the peaks Tl and the valleys T2 form the flow turns. For example, as shown in FIG 3, the flow paths F of each of the cooling passages 46A-D may include a wavy profile, with alternating peaks T l and valleys T2. In this embodiment, the peaks T l and valleys T2 are formed as smooth curves, providing smooth flow turns which prevent flow separation of the coolant at the turns Tl, T2, thereby increasing the rate of convective heat transfer. In an alternate embodiment, the flow paths F of one or more cooling passages 46A-D may have a triangular (i.e., zigzagged) profile extending along the radial direction R. [0046] To increase the turbulence in the cooling passages 46A-D, the coolant impingement openings 48A-E may be positioned on the partition walls 44A-E so as to open into flow turns T l, T2 of the cooling passages 46A-D. In the embodiment shown in FIG 3, the coolant impingement openings 48A-E are positioned at the corners of the partition walls 44-A-E, so as to open into the valleys T2 of the flow paths of the cooling passages 46A-D.

[0047] In the illustrated embodiment, as shown in FIG 5, each of the coolant impingement openings 48A-E are located substantially equidistant from the pressure side 24 and the suction side 26. In an alternate embodiment, the coolant impingement openings 48A-E may be offset from each other, i.e., have varying positions between the pressure side 24 and the suction side 26. In a further embodiment (not shown in the drawings), multiple coolant injection openings may be provided on a partition wall in the direction from the pressure side 24 to the suction side 26. In a still further embodiment, to provide improved metering of the coolant flow, the coolant impingement openings 48A-E may have different sizes along the chordal direction. For example, a chordally upstream coolant impingement opening 48 A may have a larger flow cross-section than that of a chordally downstream coolant impingement opening 48E.

[0048] In a further embodiment, turbulating features 52 may be formed on or are otherwise affixed to the inner surface 18C of the outer wall 18 (see FIG 3) within the cooling passages 46A-D. In this example, the turbulating features 52 are embodied as ribs. Alternate geometries, such as dimples or grooves may also be provided as turbulating features. The turbulating features 52 are positioned in the cooling passages 46A-D and effect a turbulation of the coolant fluid flowing therethrough so as to increase cooling provided to the outer wall 18 by coolant fluid passing through the cooling passages 46A-D [0049] The illustrated airfoil 12 further comprises a coolant fluid channel 60 that extends generally radially between the pressure side 24 and the suction side 26 and between the radially inner end 18A and the radially outer end 18B of the outer wall 18. The airfoil 12 additionally comprises a plurality of generally chordally extending outlet passages 62 formed in the outer wall 18 at the trailing edge 22. The coolant fluid channel 60 receives coolant fluid from the cooling passages 46A-D and may be configured as a single channel, as shown in FIG 1, or as multiple, radially spaced apart channels that collectively define the coolant fluid channel 60. The outlet passages 62 receive the coolant fluid from the coolant fluid channel 60 and discharge the coolant fluid from the airfoil 12 via the outlet passages 62. The coolant fluid is then mixed with the hot working gas passing through the turbine section 14. The outlet passages 62 may be located along substantially the entire radial length of the outer wall 18, or may be selectively located along the trailing edge 22 to fine tune cooling provided to specific areas.

[0050] Referring to FIGS 1 and 2, the platform assembly 16 includes an opening 68 formed therein in communication with the coolant fluid cavity 42. The opening 68 allows coolant fluid to pass from a cavity 70 (see FIG 1) formed in the platform assembly 16 into the coolant fluid cavity 42. The cavity 70 formed in the platform assembly 16 may receive coolant fluid, such as compressor discharge air, as is conventionally known in the art. [0051] The platform assembly 16 may be provided with additional openings 72, 74, 76 (see FIG 1) that supply cooling fluid to additional cavities 78, 80, 82 (see FIG 2) or sections within the hollow interior portion 28 of the outer wall 18 of the blade 12. Coolant fluid is provided from the cavity 70 in the platform assembly 16 into the cavities 78, 80, 82 to provide additional cooling to the blade 12, as will be apparent to those skilled in the art.

[0052] During operation, coolant fluid is provided to the cavity 70 in the platform assembly 16 in any known manner, as will be apparent to those skilled in the art. The cooling fluid passes into the coolant fluid cavity 42 and the additional cavities 78, 80, 82 formed in the blade 12 from the cavity 70 in the platform assembly 16, see FIGS 1 and 2.

[0053] The coolant fluid passing into the cooling fluid cavity 42 flows radially outwardly and flows serially in the chordal direction C through the cooling passages 46A-D via the coolant impingement openings 48A-E. In this embodiment, the partition wall 44A is adjoins the coolant fluid cavity 42. The coolant fluid from the coolant fluid cavity 42 flows through the coolant impingement openings 48A on the partition wall 44A adjoining the coolant fluid cavity 42 to enter the cooling passage 46A. Upon entering the cooling passage 46A, the coolant fluid impinges on the upstream face of next chordally downstream partition wall 44B as well as on the converging inner surfaces of the pressure side 24 and the suction side 26 to provide impingement cooling to these surfaces. In the cooling passage 46A, the coolant flows both in the radially outwardly as well as radially inwardly along the profiled flow path as indicated by arrows F in FIG 3. The inventive design features a series of flow turns Tl, T2 which significantly increase the flow path of the coolant fluid within the cooling passage 46A for enhancing convective heat transfer. In at least one embodiment, the coolant fluid flowing into a cooling passage 46A though a coolant impingement opening 48A on the chordally upstream partition wall 44A is made to traverse one or more flow turns T 1 , T2 before exiting the cooling passage via the next coolant impingement opening 48B on the chordally downstream partition wall 44B adjacent to the partition wall 44A, whereby the coolant fluid enters the next cooling passage 46B. This flow pattern is repeated serially in the chordal direction C to form a composite of impingement and convection cooling passage, connected in serial arrangement before exiting the trailing edge 22 of the airfoil 12. Serial impingement along the chordal direction C provides a high pressure drop and effective metering of the coolant flow rate through the trailing edge cooling feature. In the illustrated embodiment, five partition walls 44A-E are shown. However, the number of partition walls may be varied according to the overall pressure drop and coolant flow rate desired. To take advantage of the increased cooling provided by the long flow path of the coolant, the number of impingement openings in each partition wall may be reduced, which would reduce the amount of coolant flow.

[0054] Once the coolant fluid has serially traversed the last of the cooling passages, 46D, the coolant fluid passes into the coolant fluid channel 60 via the coolant impingement openings 48E on the chordally most downstream partition wall 44E. From the coolant fluid channel 60, the coolant exits the airfoil 12 via the outlet passages 62 at the trailing edge 22.

[0055] In an example embodiment, the partition walls 44A-E are cast integrally with the outer wall 18 using a casting core. An aspect of the present invention is directed to a casting core for forming a turbine airfoil with a trailing edge cooling feature in accordance with the embodiments described above.

[0056] In FIG 6, a section of the casting core 112 is shown, that corresponds to the section of the airfoil shown in FIG 3. The core 1 12 may be made, for example, of a ceramic material and may be used for manufacturing a blade assembly or a vane assembly for a turbine engine. As will be apparent to one skilled in the art, the core 112 includes a generally solid airfoil portion 1 18. The airfoil portion 118 of the core 112 is shaped to correspond to the interior contour of the outer wall 18 of the airfoil 12 to be formed, and accordingly includes a leading edge, a trailing edge, a pressure side, a suction side, a radially inner end and a radially outer end, which correspond respectively to the leading edge 20, trailing edge 22, pressure side 24, a suction side 26, radially inner end 18A and radially outer end 1 18B of the outer wall 18 of the airfoil 12 to be formed (see FIGS 1 and 2).

[0057] Referring to FIG 6, in accordance with the concepts of the present invention, the airfoil portion 118 of the core 112 includes, toward the trailing edge, a plurality of profiled recesses 144A-E interspaced by first interstitial core elements 146A-D along the chordal direction C. During casting, the profiled recesses 144A-E of the core 112 are used to form the partition walls 44A-E, while the first interstitial core elements 146A-D form the cooling passages 46A-D of the inventive airfoil 12 (see FIGS 3 and 4). The profiled recesses 144A-E extend from the pressure side to the suction side of the airfoil portion 1 18 of the core 112 and further extend generally from the radially inner end to the radially outer end of the airfoil portion 1 18 of the core 1 12. Each first interstitial core element 146A-D extends in the radial direction R from the radially inner end to the radially outer end along a contour f comprising a series of turns tl, t2, which define the profile of the cooling passages 46A-D of the inventive airfoil 12. The first interstitial core elements 146A-D are solidly interconnected via second interstitial core elements 148A-E extending through the profiled recesses 144A-E. The second interstitial core elements 148A-E of the core 112 form the coolant impingement passages 148A-E of the inventive airfoil 12. As shown, the second interstitial core elements 148A-B through adjacent profiled recesses 144A-B are staggered along the radial direction R.

[0058] In the shown embodiment, the second interstitial core elements 148A-B through adjacent first and second profiled recesses 144A-B are staggered in the radial direction R in a manner that the first interstitial core element 148A which is located between the adjacent first and second profiled recesses 144A-B exhibits at least one turn tl, t2 between a second interstitial core element 148 A through the first profiled recess 144A and a second interstitial core element 148B through the second profiled recess 144B which is radially most proximate to the second interstitial core element 148A through the first profiled recess 144A. This ensures that in the airfoil 12 formed by the core 1 12, the radial staggering between coolant impingement openings 48A-E on adjacent partition walls 44A-E is such as to include at least one flow turn T l, T2 between the nearest coolant impingement openings 48A-E on adjacent partition walls 44A-E (see FIGS 3 and 4).

[0059] In the embodiment, the second interstitial core elements 148A-E are connected to turns tl, t2 of the first interstitial core elements 146A-D. In the resulting airfoil 12 (see FIGS 3 and 4), the coolant impingement openings 48A-E open into flow turns in the cooling passages 46A-D. [0060] In the present example, at least one of the first interstitial core elements 146A-D has a periodic profile, herein a wavy profile, along the radial direction R, defined by alternating peaks tl and valleys t2 spaced apart in the radial direction R. The peaks tl and valleys t2 form the corresponding flow turns Tl, T2 in the corresponding cooling passages 46A-D of the manufactured airfoil 12. To provide smooth flow turning in the cooling passages 46A-D, the turns tl, t2 of the first interstitial core elements 146A-D comprise smooth curves.

[0061] Additionally, in the illustrated embodiment, the core 112 comprises a radially extending core portion 160 adjacent to and downstream of the last recess 144E. Further downstream, a plurality of chordally extending interstitial core elements 162, radially interspaced by gaps 163, connect the core portion 160 to the trailing edge. The core portion 160 forms the coolant fluid channel 60 while the interstitial core elements 162 form the outlet passages 62 of the manufactured airfoil 12 (see FIGS 3 and 4). [0062] During manufacture, a substrate material (not shown) may be cast around the casting core 112. The solidified cast material becomes the substrate of the airfoil 12. The casting core 1 12 may then be removed by any of several methods known to those of ordinary skill in the art. What remains once the casting core 112 is removed is the outer wall 18 delimiting the hollow airfoil interior 28 with the trailing edge cooling feature 40 integrally formed with the outer wall 18 within the airfoil interior 28.

[0063] While specific embodiments have been described in detail, those with ordinary skill in the art will appreciate that various modifications and alternative to those details could be developed in light of the overall teachings of the disclosure. Accordingly, the particular arrangements disclosed are meant to be illustrative only and not limiting as to the scope of the invention, which is to be given the full breadth of the appended claims, and any and all equivalents thereof.