EDGE

BACKGROUND 1. Field [0001] This invention relates generally to an airfoil in a turbine engine, and in particular, to a leading edge impingement cooling feature incorporated in a turbine airfoil.

2. Description of the Related Art

[0002] In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor section and then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., turbine blades, where the turbine blades extract energy from the hot combustion gases for providing output power. Since the airfoils, i.e., vanes and turbine blades, are directly exposed to the hot combustion gases, they are typically provided with internal cooling channels that conduct a cooling fluid, such as compressor bleed air, through the airfoil. [0003] One type of airfoil extends from a radially inner platform at a root end to a radially outer portion of the airfoil, and includes opposite pressure and suction sidewalls extending span-wise along a radial direction and extending axially from a leading edge to a trailing edge of the airfoil. The cooling channels extend inside the airfoil between the pressure and suction sidewalls and may conduct the cooling fluid in a radial direction through the airfoil. The cooling channels remove heat from the pressure sidewall and the suction sidewall and thereby avoid overheating of these parts. SUMMARY

[0004] Briefly, aspects of the present invention provide a turbine airfoil having an offset impingement cooling feature at the leading edge.

[0005] According to a first aspect of the invention, a turbine airfoil comprises an outer wall delimiting an airfoil interior. The outer wall extends span-wise in a radial direction and includes a pressure side and a suction side joined at a leading edge and at a trailing edge. The airfoil comprises first and second radial cavities formed in the airfoil interior. The first radial cavity is adjacent to the leading edge. The second radial cavity is adjacent to the first radial cavity. A partition wall extends from the pressure side to the suction side, separating the first and second radial cavities along a radial extent. A plurality of impingement openings are formed through the partition wall through which coolant from the second radial cavity is ejected into the first radial cavity. Each of the impingement openings is respectively configured to direct the coolant ejected into the first radial cavity to impinge on one of the pressure side or the suction side. [0006] According to a second aspect of the invention, a turbine airfoil comprises an outer wall delimiting an airfoil interior. The outer wall extends span-wise in a radial direction and includes a pressure side and a suction side joined at a leading edge and at a trailing edge. A chordal axis is defined extending centrally between the pressure side and the suction side. The airfoil comprises first and second radial cavities formed in the airfoil interior. The first radial cavity is adjacent to the leading edge. The second radial cavity is adjacent to the first radial cavity. A partition wall extends from the pressure side to the suction side, separating the first and second radial cavities along a radial extent. First and second rows of impingement openings are formed through the partition wall through which coolant from the second radial cavity is ejected into the first radial cavity. Each row of the impingement openings is arranged in a radial direction along the partition wall. In relation to the chordal axis, the impingement openings in the first row are angled toward the pressure side and the impingement openings in second row are angled toward the suction side. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] 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.

[0008] FIG 1 is a cross-sectional view through a turbine airfoil, illustrating impingement cooling of the leading edge;

[0009] FIG 2 is a cross-sectional view a portion of a turbine airfoil illustrating an offset impingement cooling feature according to an embodiment of the invention;

[0010] FIG 3 is a sectional view along the plane A-A in FIG 2 showing a radial arrangement of impingement cooling holes according to a first embodiment; [0011] FIG 4 is a sectional view along the plane A-A in FIG 2 showing a radial arrangement of impingement cooling holes according to a second embodiment; and

[0012] FIG 5 is an internal perspective view illustrating mini-grooves on an inner surface of a leading edge cavity of a turbine airfoil according to an exemplary embodiment. DETAILED DESCRIPTION

[0013] 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.

[0014] Aspects of the present invention relate to an internally cooled turbine airfoil. In a gas turbine engine, coolant supplied to the internal cooling passages in a turbine airfoil may comprise air diverted from a compressor section. 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. Embodiments of the present invention provide a cooling feature for a leading edge of an airfoil that provides increased convective heat transfer toward the leading edge, while requiring significantly lesser volume of coolant.

[0015] FIG 1 illustrates a cross-sectional view of a turbine airfoil 10, such as a turbine blade. As shown, the turbine airfoil 10 may comprise an outer wall 12 delimiting a generally hollow airfoil interior 11. The outer wall 12 is formed by a pressure side wall, referred to as pressure side 14 and a suction side wall, referred to as suction side 16. In the shown example, the pressure side 14 has a generally concave shape while the suction side 16 has a generally convex shape. The pressure side 14 and the suction side 16 are joined at a leading edge 18 and at a trailing edge 20. A chordal axis 30 may be defined as extending centrally between the pressure side 14 and the suction side 16. The outer wall 12, including the pressure and suction sides 14 and 16, extends span-wise (perpendicular to the plane of the drawing sheet) along a radial direction of the turbine engine.

[0016] The airfoil interior 11 may receive a coolant, such as air from a compressor section (not shown), via one or more coolant supply passages through a root portion of the airfoil 10 (not shown). A plurality of partition walls 22a-g are positioned in the interior portion 11, being spaced apart chordally, i.e., along the chordal axis 30. In the shown example, the partition walls 22a-g extend radially, and further extend across the chordal axis 30 connecting the pressure side 14 and the suction side 16 to define radial cavities 24a-h therebetween. The radial cavities 24a-h extend along the radial direction, being separated along a radial extent by a respective intervening partition wall 22a-g. The coolant traverses through internal cooling passages defined by the radial cavities 24a-h and exits the airfoil via rows of exhaust orifices 26a-f arranged along the span-wise or radial direction. In the shown implementation, the rows of exhaust orifices 26a-c are arranged at the leading edge 18, the rows of exhaust orifices 26d and 26e are respectively arranged at the pressure side 14 and the suction side 16 close to the leading edge 18 and the row of exhaust orifices 26f are arranged at the trailing edge 20. Exhaust orifices may be provided at additional locations along the outer wall 12 and/or at the radial tip of the airfoil 10.

[0017] In the implementation shown in FIG 1, a first radial cavity 24a is positioned adjacent to the leading edge 18 and a second radial cavity 24b is positioned adjacent to the first radial cavity 24a, such that the first radial cavity 24a is located between the second radial cavity 24b and the leading edge 18. The radial cavities 24a and 24b are separated along a radial extent by a partition wall 22a. In this case, the cooling of the leading edge 18 is accomplished by coolant impingement through a single row of crossover holes or impingement openings 28 formed on the partition wall 22a. The impingement openings 28 are positioned generally centrally between the pressure side 14 and suction side 16 and are directed toward the leading edge 18. This arrangement causes the coolant ejected from the second radial cavity 24b into the first radial cavity 24a to impinge at the leading edge 18, providing internal convection cooling at the leading edge 18. Post impingement, a portion of the coolant exits the airfoil 10 through the rows of exhaust orifices 26a-c located at the leading edge 18, providing film cooling at the outer surface of the leading edge 18. The exhaust openings 26a-c may be designed, for example as rows of showerhead holes. The remaining portion of the post impingement coolant travels internally along the outer wall 12 toward the exhaust orifices 26d and 26e located respectively at the pressure side 14 and suction side 16 adjoining the radial cavity 24a, to provide convection cooling to these parts. The exhaust orifices 26d-e may be designed as gill-holes, i.e., angled away from the leading edge 18, such that the coolant discharged from the exhaust orifices 26d-e provides film cooling on the external surface of the pressure side 14 and suction side 16 near the leading edge 18. The above described design is particularly effective in cooling leading edges having smaller diameter of curvature.

[0018] In many advanced turbine designs, the turbine airfoil leading edge may be made larger to provide an increased internal cooling surface area in relation to the external heated surface area. Furthermore, for achieving increased turbine efficiency, lower coolant flows may be desired. To meet the requirement of lower coolant flows, the gill holes 26d-e may be eliminated. The present inventors have recognized that under these circumstances, the above-described single row of impingement openings 28 may result in non-effective cooling on the pressure side 14 and the suction side 16 of the leading edge cavity 24a. In particular, it has been seen that impingement of the coolant directly at the leading edge causes most of the coolant to exit the airfoil through the showerhead holes at the leading edge. Since the coolant flow is already low and the internal surface area is higher in a large diameter leading edge, such an arrangement may lead to insufficient internal convention cooling at the leading edge.

[0019] FIG 2 shows a cross-sectional view of a portion of a turbine airfoil 10, illustrating a leading edge cooling feature in accordance with an embodiment of the present invention. For the sake of simplicity, the description of like elements with respect to FIG 1 will not be reiterated. In the illustrated embodiment, the airfoil 10 is a turbine blade for a gas turbine engine. It should however be noted that aspects of the invention could additionally be incorporated into stationary vanes in a gas turbine engine. As shown, the airfoil 10 has a rounded leading edge 18 having a larger diameter in comparison to that shown in FIG 1. A plurality of radial cavities defining internal cooling passages may be formed in the airfoil interior 11 , of which only radial cavities 24a and 24b are illustrated in FIG 2. Coolant fluid, for example, air from a compressor section, may traverse through the cooling passages defined by the radial cavities and exit the airfoil 10 via exhaust orifices, of which only the exhaust orifices 26a,26b and 26c at the leading edge 18 are illustrated. The exhaust orifices 26a-c may be embodied as multiple rows of showerhead holes. In the illustrated embodiment, to meet a low coolant flow requirement, gill holes are eliminated. That is, the outer wall 12 adjacent to the first radial cavity 24a is formed without exhaust openings at the pressure side 14 or at the suction side 16. [0020] As shown in FIG 2, instead of a single row of impingement openings aimed at the leading edge, first and second rows of impingement openings 28a, 28b are formed through the partition wall 22a, through which coolant from the second radial cavity 24b is ejected into the first radial cavity 24a adjacent to the leading edge 18. The impingement openings 28a, 28b are respectively configured to direct the coolant ejected into the first radial cavity 24a to impinge directly on the pressure side 14 and the suction side 16. As shown, in relation to the chordal axis 30, the impingement openings 28a in the first row are angled toward the pressure side 14 and the impingement openings 28b in second row are angled toward the suction side 16. In one embodiment, each impingement opening 28a in the first row may be configured to have a respective flow axis 32a that intersects with the pressure side 14. Likewise, each impingement opening 28b in the second row may be configured to have a respective flow axis 32b that intersects with the suction side 16. In contrast to FIG 1, in the embodiment of FIG 2, the first and second rows of impingement openings 28a and 28b are offset with respect to the chordal axis 30. In particular, the first row of impingement openings 28a is positioned closer to the pressure side 14 than the suction side 16, while the second row of impingement openings 28b is positioned closer to the suction side 16 than the pressure side 14.

[0021] The angled configuration of the impingement openings 28a-b ensures that the impinged coolant is targeted on both the pressure side 14 as well as the suction side 16 of the leading edge radial cavity 24a. Post impingement, the coolant travels along the inner surface 13 of the pressure side 14 and the suction side 16 toward the exhaust orifices 26a- c at the leading edge 18 (indicated by arrows 38), from where it is discharged into the hot gas path. A relatively large surface area of the inner surface 13 is thereby traversed by the coolant for aiding internal convective cooling. The coolant exiting from the exhaust orifices 26a-c may further provide film cooling on the external surface of the leading edge (indicated by arrows 40).

[0022] Each row of the impingement openings 28a, 28b may be arranged in a radial direction along the partition wall 22a. In one embodiment, as shown in FIG 3, the impingement openings 28a in the first row may be staggered in a radial direction in relation to the impingement openings 28b in the second row. Such an arrangement may provide manufacturing advantages, for example if the airfoil is formed by casting using a ceramic casting core. In this case, the staggered arrangement provides a zigzagged shaped core tooling parting line (indicated by dashed line 34 in FIG 3). This ensures that when the die used for forming the casting core is opened, the impingement openings/holes are not cut into. In other embodiments, alternate manufacturing techniques, for example, additive manufacturing techniques such as selective laser melting (SLM), may be used for forming the impingement openings. The use of additive manufacturing techniques may also allow the adjacent rows of impingement openings 28a, 28b to be arranged in a non-staggered (i.e., radially aligned) manner. For example, in an alternate embodiment as shown in FIG 4 corresponding impingement openings 28a, 28b in the first row and the second row respectively may be positioned at equal radial heights along the partition wall 22a.

[0023] In a further embodiment, as shown in FIG 5, the inner surface 13 of the outer wall 12 adjoining the radial cavity 24a comprises grooves 36a, 36b to enhance cooling surface area for better heat transfer. The grooves 36a, 36b may be positioned at least at the locations where the coolant impinges at the pressure side 14 and suction side 16 respectively. In the illustrated embodiment, a first row of grooves 36a is provided on the inner surface 13 at the pressure side 14, while a second row of grooves 36b is provided on the inner surface 13 at the suction side 16. The grooves 36a, 36b in each row may extend parallel to each other, with each individual groove 36a-b extending at a fixed radius along the pressure side 14 or suction side 16 toward the exhaust orifices 26a-c at the leading edge 18.

[0024] In a still further embodiment, each groove 36a, 36b may be sized such that a radial dimension of each groove is equal to or smaller than a diameter of an impingement opening 26a, 26b from which coolant impinges on that groove 36a, 36b. Each groove may thereby have a size equal to or smaller than the jet size of the impinging jet. It may thereby be ensured that a single impingement jet is able to fill multiple grooves, and not contained in a single groove, thereby enhancing convective heat transfer area. As a further feature, the impingement openings 28a-b in each of the first row and second row may be oriented at varying angles in relation to the radial direction, to form a divergent pattern of coolant jets impinging on the pressure side 14 and/or the suction side 16. Such a feature may also ensure that an impingement jet is able to fill a large number of grooves, to provide still enhanced cooling. [0025] 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.