BACKGROUND 1. Field

[0001] The present invention is directed generally to turbine airfoils, and more particularly to turbine airfoils having internal cooling channels for conducting a coolant through the 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., turbine vanes and blades, are directly exposed to the hot combustion gases, they are typically provided with internal cooling channels that conduct a coolant, 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 coolant in alternating radial directions 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 internal cooling channel with a profiled flow blocking structure for enhanced near wall cooling. [0005] According to a first aspect of the invention, an airfoil for a turbine engine is provided. The airfoil comprises an airfoil section comprising an outer wall extending span-wise along a radial direction of the turbine engine and being formed of a pressure side and a suction side joined at a leading edge and a trailing edge. A flow blocking structure is positioned in an interior of the airfoil section for displacing a coolant flow toward the pressure side and the suction side. The flow blocking structure comprises a hollow body defining a radial inactive cavity. The hollow body comprises a first side wall spaced from the pressure side to define a first primary cooling passage for conducting coolant radially along the pressure side. The hollow body further comprises a second side wall opposite to the first side wall and spaced from the suction side to define a second primary cooling passage for conducting coolant radially along the suction side. The first side wall and/or second side wall is profiled along a radial extent, to locally decrease a flow cross-section of the first and/or the second primary cooling passage respectively, and increase heat transfer between the coolant and outer wall.

[0006] According to a second aspect of the invention, a machined casting is provided. The machined casting comprises an airfoil section having an outer wall delimiting a generally hollow airfoil interior. The outer wall extends span-wise and is formed of a pressure side and a suction side joined at a leading edge and a trailing edge. A flow blocking structure is formed monolithically with the airfoil section and is positioned in the airfoil interior. The flow blocking structure comprises a hollow body defining an inactive cavity extending in a span-wise direction. The hollow body comprises a first side wall spaced from the pressure side to define a first cooling channel. The hollow body also comprises a second side wall opposite to the first side wall and spaced from the suction side to define a second cooling channel. The first side wall and/or second side wall is profiled along a length in the span-wise direction, to locally decrease a cross-sectional area of the first and/or the second cooling channel respectively.

[0007] Those skilled in the art may apply the respective features of the present invention jointly or severally in any combination or sub-combination.

BRIEF DESCRIPTION OF THE DRAWINGS [0008] The teachings of the present disclosure can be readily understood by considering the following detailed description in conjunction with the accompanying drawings, in which:

[0009] FIG 1 is a perspective view of an example of a turbine airfoil according to one embodiment; [0010] FIG 2 is a cross-sectional view through the turbine airfoil along the section II- II of FIG 1;

[0011] FIG 3A, 3B and 3C are sectional views along the sections A-A, B-B and C-C respectively in FIG 2, illustrating an exemplary embodiment of the present invention; and

[0012] FIG 4 is a flow diagram illustrating an exemplary serpentine flow scheme through the airfoil.

[0013] To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures.

DETAILED DESCRIPTION

[0014] Although various embodiments that incorporate the teachings of the present disclosure have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The scope of the disclosure is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The disclosure encompasses other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of "including," "comprising," or "having" and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items.

[0015] 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 often comprises air diverted from a compressor section. Thermal efficiency of a gas turbine engine may be increased by lowering the volume of coolant diverted from the compressor section to the internal cooling channels inside the turbine airfoils. This may be achieved by providing at least one flow blocking structure in the airfoil interior to reduce the cross-sectional flow area of one or more radial internal cooling channels. A flow blocking structure may be embodied as a hollow elongated body defining a radial cavity therewithin. The cavity may be inactive, that is to say that there may be no coolant flow through the cavity. The inactive cavity serves to occupy space in the airfoil interior, to displace the radially flowing coolant toward near- wall passages adjacent to the hot pressure and suction side walls. This provides more effective cooling of the hot outer wall of the airfoil in comparison to a typical two-wall turbine airfoil where a significant portion of the radial coolant flow remains toward the center of the flow cross-section between the pressure and suction side walls, and is hence underutilized for convective cooling. One of the objectives of the flow blocking structure is to provide a narrower near-wall flow cross-sectional area, whereby higher coolant velocities and thereby higher convective heat transfer may be achieved at reduced coolant flow rates. The present inventor has recognized that, especially for airfoils manufactured by casting, near-wall passages may desirably be of a sufficient size to ensure a strong ceramic core for casting, which may limit the amount by which flow cross-section of the near-wall passages can be reduced and thus limit the coolant velocities that may be achieved. Embodiments of the present invention address at least some of the above issues in connection with a flow blocking structure. [0016] Referring now to FIG 1, a turbine airfoil 10 is illustrated according to one embodiment. As illustrated, 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. The airfoil 10 may include a generally elongated hollow airfoil section 12 formed from an outer wall 14 adapted for use, for example, in a high pressure stage of an axial flow gas turbine engine. The outer wall 14 extends span-wise along a radial direction R of the turbine engine and includes a generally concave shaped pressure side 16 and a generally convex shaped suction side 18. The pressure side 16 and the suction side 18 are joined at a leading edge 20 and at a trailing edge 22. As illustrated, the generally elongated hollow airfoil section 12 may be coupled to a root 56 at a platform 58. The root 56 may couple the turbine airfoil 10 to a disc (not shown) of the turbine engine. The generally hollow airfoil section 12 is delimited in the radial direction by a radially outer end face or airfoil tip 52 and a radially inner end face 54 coupled to the platform 58. In other embodiments, the airfoil 10 may be a stationary turbine vane with a radially inner end face coupled to the inner diameter of the turbine section of the turbine engine and a radially outer end face coupled to the outer diameter of the turbine section of the turbine engine.

[0017] Referring to FIG 2, a chordal axis 30 is defined extending generally centrally between the pressure side 16 and the suction side 18 from the leading edge 20 to the trailing edge 22. As illustrated, the generally hollow elongated airfoil section 12 comprises an interior portion 11 which may receive a coolant, such as air from a compressor section (not shown), via one or more coolant supply passages (not shown) through the root 56. A plurality of partition walls 24 are positioned in the interior portion 11 spaced apart chordally, i.e., along the chordal axis 30. The partition walls 24 extend radially, and further extend across the chordal axis 30 along a substantially straight profile connecting the pressure side 16 and the suction side 18 to define radial cavities or flow passes 41, 42, 43, 44, 45, 46, 47. The coolant traverses through the flow passes 41- 47 and exits the airfoil section 12 via exhaust orifices 27 and 29 positioned along the leading edge 20 and the trailing edge 22 respectively. The exhaust orifices 27 provide film cooling along the leading edge 20 (see FIG 1). Although not shown in the drawings, film cooling orifices may be provided at multiple locations, including anywhere on the pressure side 16, suction side 18, leading edge 20 and the airfoil tip 52.

[0018] According to the illustrated embodiment, one or more flow blocking structures 26 may be provided in the interior portion 11 of the airfoil section 12. Each flow blocking structure 26 may be positioned between a pair of adjacent partition walls 24 and may include an elongated hollow body 28 extending lengthwise along the radial direction R. In the illustrated embodiment, each hollow body 28 defines a respective inactive cavity Tl, T2 within. An inactive cavity may not have any coolant flow through the inactive cavity. For example, each of the inactive cavities Tl, T2 may be a dead space that does not include an active flow of fluids, but serve to reduce the cross-sectional area of the radial flow of the coolant and further to displace the coolant toward the pressure side 16 and the suction side 18. Each of the inactive cavities Tl, T2 extends radially and is isolated from the outer wall 14. The generally hollow construction of the flow blocking structure 26 incorporating an inactive cavity Tl, T2 within may provide reduced thermal stresses as compared to a solid body construction and may further reduce centrifugal loads in case of a rotating blade. The hollow body 28 may extend across the chordal axis 30, being spaced from the pressure 16 and suction 18 sides. The hollow body 28 includes first 82 and second 84 opposite side walls that respectively face the pressure 16 and suction 18 sides. In the illustrated embodiment, connector ribs 32, 34 may extend respectively from the first 82 and second 84 opposite side walls to the pressure side 16 and suction side 18. The pair of connector ribs 32, 34 thus connect the hollow body 28 to the pressure 16 and suction 18 sides along a radial extent.

[0019] The first side wall 82 is spaced from the pressure side 16 to define a first primary cooling passage 72 for conducting coolant radially along the pressure side 16. The second side wall 84 is spaced from the suction side 18 to define a second primary cooling passage 74 for conducting coolant radially along the suction side 18. The primary cooling passages 72, 74 provide near wall cooling along the pressure side 16 and the suction side 18. In the illustrated embodiment, the first side wall 82 is generally parallel to the pressure side 16 of the outer wall 14 and the second side wall 84 is generally parallel to the suction side 18 of the outer wall 14. The hollow body 28 may further comprise forward 86 and aft 88 end walls that may extend between the first 82 and second 84 side walls and may be spaced along the chordal axis 30. In alternate embodiments, the hollow body 28 may have, for example, a triangular, circular, elliptical, oval, polygonal, or any other shape or outer contour essentially comprising a first side wall facing the pressure side and defining a first primary cooling passage therebetween, and a second side wall facing the suction side and defining a second primary cooling passage therebetween. The first 72 and second 74 primary cooling passages may be connected along a radial extent via a secondary cooling passage 76 to form a radial flow pass 43,44,45,46 with a generally C-shaped flow cross-section. In the illustrated embodiment, each secondary cooling passage 76 is formed between a respective end wall 86, 88 of the hollow body 28 and one of the adjacent partition walls 24.

[0020] As shown in FIG 2, a pair of adjacent radial flow passes 43, 44 or 45,46 are defined on opposite sides of each flow blocking structure 26 along the chordal axis 30. For example, a first pair of adjacent radial flow passes 43, 44 is defined on chordally opposite sides of a first flow blocking structure 26, while a second pair of adjacent radial flow passes 45, 46 is defined on chordally opposite sides of a second flow blocking structure 26. As noted above, each of the radial flow passes 43-46 comprises a generally C-shaped flow cross-section formed by a first primary cooling passage 72 adjacent to the pressure side 16, a second primary cooling passage 74 adjacent to the suction side 18, and a secondary cooling passage 76 connecting the first and second primary cooling passages 72, 74. The provision of secondary cooling passages 76 connecting the primary cooling passages 72, 74 provides reduced thermal stresses caused by differential expansion of the relatively colder and hotter pars of the airfoil 10. As illustrated in FIG 2, the radial flow passes of each pair 43,44 or 45,46 have symmetrically opposed C-shaped flow cross- sections. That is, the flow cross-section of the radial flow pass 44 corresponds to a mirror image of the flow cross-section of the radial flow pass 43, with reference to a mirror axis generally perpendicular to the chordal axis 30. The same description holds for the second pair of radial flow passes 45, 46. It should be noted that the term "symmetrically opposed" in this context is not meant to be limited to an exact dimensional symmetry of the flow cross-sections, which often cannot be achieved especially in highly contoured airfoils. Instead, the term "symmetrically opposed", as used herein, refers to symmetrically opposed relative geometries of the elements that form the flow cross- sections (i.e., the primary cooling passages 72, 74 and the secondary cooling passage 76 in this example). As described in detail hereinafter referring to FIG 4, the radial flow passes of each pair 43,44 or 45,46 conduct coolant in opposite radial directions and are fluidically connected in series via a chordal flow passage 50a, 50c (i.e., parallel to the chordal axis 30) defined in the interior portion 11 by a gap 51 (see FIGS 3A-C) between the flow blocking structure 26 and a radial end face 52 of the airfoil section 12, to form a serpentine cooling path 60a, 60b.

[0021] Referring to FIGS 3A-C, each flow blocking structure 26 including the hollow body 28 extends in the radial direction R with one of the radial ends 38 terminating within the interior portion 11 of the hollow airfoil section 12 short of a radial end face 52, in this case the airfoil tip, to define a gap 51. The other radial end 36 (shown schematically in FIG 4) may be located, for example, at the root portion 56 of the airfoil 10. At least one of the radial ends of each flow blocking structure 26 may be closed or capped to ensure that there is no fluid flow through the inactive cavities Tl, T2. In the illustrated embodiment, the radial end 38 is capped via a tip cap 37.

[0022] The present inventor has devised a structure wherein one or both of the side walls 82 and 84 facing the hot outer wall 14 are profiled along a radial extent to locally decrease a flow cross-section of the first 72 and/or the second 74 primary cooling passages respectively, to enhance heat transfer between the radially flowing coolant 62 and the outer wall 14. In this example, both of the side walls 82, 84 of each of the flow blocking structures 26 are corrugated in the radial direction R, defining a periodic profile having local peaks 64 and troughs 66. The peaks 64 define local minimum flow cross- sectional areas of the primary cooling passages 72, 74 while the troughs 66 define local maximum flow cross-sectional areas of the primary cooling passages 72, 74. The peaks 64 and troughs 66 may extend perpendicular to the radial direction R, along the first 82 and/or second 84 side walls of the flow blocking structure 26. [0023] In the shown example, the local flow cross-sectional area at the peaks 64 is reduced in comparison to a baseline configuration with straight radial side walls, schematically designated as 90. The constriction of the flow cross-section at the peaks 64 locally accelerates the coolant 62 flow and increases the overall coolant 62 velocity in the radial direction R, thereby enhancing convective heat transfer coefficient between the coolant 62 and the outer wall 14. At the same time, by providing an increased flow- cross-section area at the troughs 66, it is ensured that the flow blocking structure 26 provides a sufficient ceramic volume in the near wall (primary) cooling passages for sufficient core strength during casting of the airfoil. Another technical effect achieved by the illustrated embodiment is that the periodic profile of the side walls 82 and 84 creates a periodic shear jet adjacent to the hot outer wall 14 which interrupts boundary layer thickness growth along the outer wall 14 at the pressure side 16 and the suction side 18. This may be attributed to the fact that the periodic variation of the flow cross-section restarts the boundary layer periodically, thereby preventing its growth. The reduction of boundary layer growth leads to increased heat transfer between the coolant 62 and the outer wall 14. In the illustrated embodiment, each of the side walls 82 and 84 is shown to have a wave-like profile, with the peaks 64 and troughs 66 providing smooth flow turns, to prevent a high pressure drop in the serpentine flow path. A wave-like profile with gentle peaks and troughs may also reduce stress concentrations in the ceramic core used for manufacture of the airfoil. Notwithstanding, other periodic profile shapes, e.g. triangular, square, etc., may be applied as apparent to one skilled in the art.

[0024] As illustrated, the periodic profile is associated with an amplitude PA and a pitch or periodicity Pp. A higher amplitude PA provides increased flow constriction to create higher coolant velocities, while a higher pitch P _P may provide further reduction in boundary layer growth. In one embodiment, the pitch or periodicity Pp of the profile may be varied as a function of local heat transfer requirement. Additionally or alternately, the amplitude PA of the profile may be varied as a function of local heat transfer requirement. In one embodiment, the corrugation may extend along the entire radial length of the first 82 and/or second 84 side wall. In alternate embodiments, the corrugations may be localized to only a portion of the entire radial extent of the first side wall 82 and/or the second side wall 84 as a function of local heat transfer requirement. Furthermore, it may be apparent to one skilled in the art to use differing profile shapes and parameters (pitch, amplitude, location etc.) between the first 82 and second 84 side walls, as also between different flow blocking structures 26 of the airfoil 10, as function of the amount of cooling required.

[0025] Referring now to FIG 4, a non-limiting example of a cooling scheme is illustrated. The illustrated cooling scheme involves two independent and oppositely directed serpentine cooling paths, namely a first serpentine cooling path 60a and a second serpentine cooling path 60b. The first serpentine cooling path 60a begins at the radial flow pass 44 that receives a coolant from a respective coolant supply that may be located at the root 56 of the turbine airfoil 10. The coolant then flows in alternating radial directions through the radial flow passes 44, 43 and 42. The radial flow passes 44, 43 and 42 are connected in series via respective chordal flow passages 50a, 50b to form a serpentine path 60a extending along an aft-to-forward direction along the chordal axis 30. From the radial flow pass 42, the coolant fluid enters a radial cavity 41 via impingement openings 23, and then discharged into the hot gas path via exhaust orifices 27 on the outer wall which collectively form a shower head for cooling the leading edge 20 of the turbine airfoil 10. The second serpentine cooling path 60b begins at the radial flow pass 45 that receives a coolant from a respective coolant supply that may be located at the root 56 of the turbine airfoil 10. The coolant then flows in alternating radial directions through the radial flow passes 45, 46 and 47. The radial flow passes 45, 46 and 47 are connected in series via respective chordal flow passages 50c, 50d to form a serpentine path 60b extending along a forward-to-aft direction along the chordal axis 30. The radial cavity 47 may incorporate trailing edge cooling features 49 (FIG 2), as known to one skilled in the art, for example, comprising turbulators, or pin fins, or combinations thereof, before being discharged into the hot gas path via exhaust orifices 29 located along the trailing edge 22.

[0026] The gap 51 in the interior portion 11 of the hollow airfoil section 12, in cooperation with the symmetrically opposed flow cross-sections of the pair of adjacent radial flow passes 43, 44 or 45, 46, ensures a uniform flow turn at the chordal passages 50a, 50c from an upstream radial flow pass 44 or 45 to a respective downstream radial flow pass 43 or 46 in the respective serpentine cooling path 60a, 60b. The gap 51 also reduces mechanical stresses experienced by the flow blocking structure 26 due to differential thermal expansion with respect to the relatively hot pressure and suction side walls 16 and 18, and further provides convective shelf cooling of the radial end face 52 of the airfoil section 12. The chordal flow passages 50b and 50d may be located for example, in the interior portion 11 of the hollow airfoil section 12, or at the platform 58.

[0027] Embodiments of the present invention enhance convective heat transfer in the serpentine cooling paths 60a, 60b, by locally reducing the flow cross-sectional area of the primary cooling (near wall) passages 72, 74 of the radial flow passes 43-46 and accelerating coolant flow. Furthermore, the wave-like profile prevents excessive pressure drop of the coolant in the serpentine path before being delivered to the downstream impingement features toward the leading and trailing edges. The convective heat transfer provided by the present invention can be further facilitated by additional features known to one skilled in the art. For example, turbulator features, such as turbulator ribs, may be formed on the inner surface of the hot outer wall 14 in the radial flow passes of the serpentine cooling paths. For example, the turbulator ribs can be configured to prevent overcooling at the upstream end of the serpentine cooling paths. In particular, the number and size of the turbulator ribs can be varied along the cooling path, such as by providing an increased turbulator count, and providing larger turbulator ribs, in the downstream direction to increase the heat transfer effect of the turbulator ribs in the downstream direction of the cooling path as the cooling air warms, to thereby enable the heated cooling air to remove an adequate amount of heat from the outer wall in the downstream direction. Additionally, flow guides may be provided, especially in the C-shaped radial flow passes, which optimize the flow distribution in these flow passes and target wall heat transfer. Furthermore, refresher feeds may be incorporated in the serpentine cooling paths. Still further, film cooling holes may be provided on various locations along the serpentine cooling paths for local cooling. [0028] In the illustrated example, the flow blocking structure is manufactured integrally with the airfoil section by a process of casting using a ceramic casting core, followed by machining. The corrugated design ensures sufficient core material for good strength while accelerating the cooling flow in the near-wall passages. In alternate embodiments, the flow blocking structure may be manufactured integrally with the airfoil section using any other manufacturing technique that does not require post manufacturing assembly as in the case of inserts. Other manufacturing techniques may include, for example, additive manufacturing processes such as 3-D printing. A monolithic construction allows the inventive design to be used for highly contoured airfoils, including 3-D contoured blades and vanes, where inserts may not be usable.

[0029] 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.