BACKGROUND

1. Field

[0001] The present invention relates to turbine components such as rotating blades and stationary vanes, and in particular, to a blade or vane platform cooling circuit with mate face cooling.

2. Description of the Related Art

[0001] A gas turbine engine may include a compressor section for compressing air, a combustor section for mixing the compressed air with fuel and igniting the mixture to form a hot working fluid, and a turbine section for producing power from the hot working fluid. The turbine section is usually provided with multiple alternating rows of stationary vanes and rotor blades that expand the hot working fluid to produce mechanical power. Typically, a rotor blade may include an airfoil extending span- wise from an inner diameter platform, while a stationary vane may include an airfoil extending span-wise between an inner diameter platform and an outer diameter platform. A plurality of turbine blades or vanes are positioned circumferentially adjacent to each other to form a respective row of blades or vanes. The efficiency of a gas turbine engine can be increased by passing a higher temperature gas flow into the turbine section. Turbine blades and vanes often contain cooling systems for prolonging the life of these components and reducing the likelihood of failure as a result of excessive temperatures.

[0002] In particular, it is desirable to provide cooling air to mate face gaps between circumferentially adjacent components (e.g. blades or vanes) in order to cool the mate face region of the platform and prevent hot gas from ingesting into the mate face gap and past a circumferential damper seal. Hot gas ingestion into the mate face gap may lead to degradation of the component and/or the damper seals, causing premature failure of the component before the design life is met.

[0003] It is known to provide platform mate face cooling using cooling air delivered from a platform cooling circuit. FIG. 1 illustrates a known configuration of a cooling circuit 104 for a platform 102 employed in a rotor blade. The cooling circuit 104 may have an inlet connected to an inner core of a blade airfoil 100 and includes a cooling channel 106 connected to a series of cooling holes 108 through a platform mate face 110. By exhausting the spent cooling air of the platform cooling circuit through the mate face, the cooling objectives of both the platform and the mate faces may be met with a single coolant flow source.

SUMMARY

[0004] Briefly, aspects of the present invention relate to a turbine blade or vane platform cooling circuit with improved mate face cooling.

[0005] According to an embodiment of the present invention, a turbine component is disclosed, such as a rotor blade or a stationary vane. The component includes at least one platform and an airfoil extending span-wise from the platform. The airfoil comprises a pressure side wall and a suction side wall joined at a leading edge and a trailing edge. The platform is provided with a platform cooling circuit, comprising: an inlet for receiving a coolant; a cooling channel formed within the platform in fluid communication with the inlet, the cooling channel positioned spaced from a mate face of the platform and extending along at least a portion of the length of the mate face; and a plurality of outlet holes connected to the cooling channel and opening into the mate face. At least one rib is positioned in the cooling channel, which may be configured to divide a coolant flow through the cooling channel into multiple flow channels including at least a first flow channel and a second flow channel. The first flow channel is connected to a first subset of the outlet holes, and the second flow channel is connected to a second subset of outlet holes, which is located downstream of the first subset of the outlet holes in relation to a flow direction through the cooling channel.

BRIEF DESCRIPTION OF THE DRAWINGS

[0006] 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. [0007] FIG. 1 is a schematic radial top view of a turbine blade illustrating a known configuration of a platform cooling circuit with mate face cooling;

[0008] FIG. 2 is a perspective view of a turbine blade wherein aspects of the present invention may be employed;

[0009] FIG. 3 is a perspective top view of a turbine blade, illustrating a platform cooling circuit according to one embodiment of the present invention;

[0010] FIG. 4 depicts a temperature distribution map of a turbine blade with platform cooling circuit having a split cooling channel;

[0011] FIG. 5 depicts a temperature distribution map of a turbine blade with platform cooling circuit having an un-split cooling channel; and

[0012] FIG. 6 is a schematic sectional view, looking in a direction from the pressure side to the suction side, showing a cooling channel embodied on a platform with 3D end-wall contouring.

DETAILED DESCRIPTION

[0013] In the following detailed description of the preferred embodiments, 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] The description and the drawings illustrate an example embodiment of a turbine component wherein aspects of the present invention may be incorporated. Although the component illustrated herein is a turbine rotor blade, one skilled in the art would appreciate that the underlying teachings would apply also to a turbine stationary vane.

[0015] As shown in FIG. 2, a turbine blade 1 comprises an airfoil 10 extending span-wise from a platform 6. The airfoil 10 comprises a generally concave pressure side wall 12 and a generally convex suction side wall 14 (see also FIG. 3), joined at a leading edge 16 and at a trailing edge 18. An attachment structure 3, referred to as a root portion extends radially inward from the platform 6 via a shank 5 for affixing the blade 1 to a rotor disc (not shown). The turbine blade 1 is rotatable about an engine axis 1 1. The platform 6 comprises a radially outer surface 6a that forms an inner flow path boundary for a hot gas, and a radially inner surface 6b facing a rim cavity 9. The platform is delimited along an axial direction by a forward face 26 and an aft face 28.

[0016] Along a circumferential direction, the platform 6 is delimited by a pressure side mate face 22 and a suction side mate face 24 extending radially between and typically (but not necessarily) orthogonal to the radially outer surface 6a and the radially inner surface 6b of the platform. Upon assembly of the blade row, the pressure side mate face 22 of a blade faces a suction side mate face 24 of a circumferentially adjacent blade, with a mate face gap defined therebetween. As previously stated, it is often desirable to provide cooling air to the mate face gap between circumferentially adjacent turbine components (such as blades and vanes) in order to cool the mate face region of the platform and prevent hot gas from ingesting into the mate face gap. To this end, in the shown embodiment, the pressure side mate face 22 is provided with cooling holes 70 to deliver a coolant from a platform cooling circuit 30 (an example of which is shown in FIG. 3) to the mate face gap between adjacent blades. Although not specifically shown, cooling holes may also be provided at the suction side mate face 24.

[0017] It has been recognized that in a standard platform cooling circuit with mate face cooling (such as that exemplified in FIG. 1), as the cooling flow is exhausted through the series of outlet holes located at the mate face, the amount of cooling flow is reduced down the remaining length of the cooling channel that feeds into the outlet holes. The amount of heat that can be removed from a surface through convective cooling is directly proportional to the amount of coolant moving over that surface. Therefore, as the amount of coolant is reduced over the length of the cooling channel, the heat transfer over that length is consequently reduced. Such a reduced heat transfer causes the component temperatures above the cooling channel (which is this case is the platform outer surface) to increase, leading to hot spots. The present inventor has devised a cooling mechanism that addresses at least the above-described problem and provides a targeted mate face cooling.

[0018] Referring now to FIG. 3, an example embodiment of an inventive platform cooling circuit 30 is described. The platform cooling circuit 30 comprises at least one cooling channel 40 located spaced from a mate face 22 of the platform, which in this example is the pressure side mate face. As shown, the cooling channel 40 may be located nearer to the mate face 22 than to the airfoil pressure side wall 12. The cooling channel 40 is formed entirely within the platform 6, defining a closed flow conduit between the radially outer and inner surfaces 6a and 6b respectively of the platform 6. The cooling channel 40 extends along at least a portion of the mate face 22, in this example, extending for a major portion of the axial length of the mate face 22. In the illustrated embodiment, the cooling channel 40 is part of a serpentine platform cooling passage, forming the last leg of said serpentine cooling passage. A plurality of outlet holes 70 are connected to the cooling channel 40, opening into the mate face 22. One or more dividing ribs 60 may be positioned in the cooling channel 40, for at least a portion of the length of the cooling channel 40, so as to divide or split the coolant flow through the cooling channel 40 into multiple parallel flow channels. In the non-limiting example illustrated herein, a single rib 60 is provided, which divides the coolant flow in the cooling channel 40 into first and second parallel flow channels 40a and 40b respectively. The first flow channel 40a is connected to a first subset 70a of the outlet holes 70, while the second flow channel 40b is connected to a second subset 70 of the outlet holes 70. The second subset 70b of outlet holes 70 is located downstream of the first subset 70a of outlet holes 70 in relation to a flow direction through the cooling channel 40. In further embodiments (not shown), multiple dividing ribs may be disposed in the cooling channel to split the coolant flow into three or more parallel flow channels, each supplying coolant to a separate subset of mate face cooling outlet holes.

[0019] By splitting the cooling channel into parallel flow channels, each supplying coolant to a subset of the mate face cooling outlet holes, the region of reduced heat transfer on account of reduced amount of coolant is minimized in relation to an un- split channel. The above effect is illustrated by a comparison of FIG. 4 and FIG. 5. FIG. 4 depicts a temperature distribution map 82 of a turbine blade with a serpentine platform cooling circuit with a split cooling channel supplying coolant to mate face outlet holes (as shown in FIG. 3); while FIG. 5 depicts a temperature distribution map 84 of a turbine blade with an essentially similar platform cooling circuit with the exception of not having a dividing rib in the cooling channel (i.e., with an un-split cooling channel). As shown in FIG. 5, the temperature of the platform increases uniformly moving from the pressure side wall of the airfoil toward the pressure side mate face along the platform cooling circuit. In particular, towards the downstream end of the platform cooling circuit, the temperature rapidly increases to produce hot spots 92 in a region 90 of reduced heat transfer where the mate face cooling air is exhausted. In comparison, as seen in FIG. 4, a turbine blade platform with a split cooling channel exhibits a significantly lower temperature in the same area 90 which corresponds to the region of reduced heat transfer in a turbine blade platform with an un-split cooling channel. The above effect is facilitated, at least in part, by the constriction of flow by the splitting of the cooling channel, which leads to increased velocity of the coolant proceeding towards the outlet holes 70a, 70b, thereby increasing the convective heat transfer coefficient near the platform mate face as well as directly beneath the outer platform surface 6a.

[0020] Referring back to FIG. 3, the airfoil 10 of the turbine blade 1 may be have a number of radially extending internal cavities 50, defined between the pressure side wall 12 and the suction side wall 14 by partitioning ribs 52. The internal cavities 50 may form one or more airfoil cooling circuits circulating cooling air, which may be received, for example, from a compressor section of the engine via coolant supply passages (not shown) formed in the blade root. In one embodiment, to minimize coolant consumption, the platform cooling circuit 30 may be integrated with an airfoil cooling circuit, utilizing spent coolant from the airfoil cooling circuit. To this end, the inlet 32 of the platform cooling channel 30 may be connected to an internal cavity 50 of the airfoil to receive the coolant K which was used in cooling the airfoil 10. In the present example, the inlet 32 is connected to an internal cavity 50a in an aft portion of the airfoil 10. The connection may be located radially inboard of the outer surface 6a of the platform 6. One skilled in the art would appreciate that the inlet 32 may be connected to any of the other internal cavities 50. In an alternate embodiment, the inlet 32 of the platform cooling circuit 30 may receive coolant supply from the rim cavity 9 underneath the platform 6. In still other embodiments, especially in case of stationary vanes, the platform cooling circuit inlet may receive coolant supply from a coolant source radially inboard of an ID vane platform or radially outboard of an OD vane platform. In the illustrated embodiment, to provide increased mate face cooling, the platform cooling circuit 30 is devoid of any other outlet holes through the platform surface, such that the entirety of the coolant K in the platform cooling circuit 30 is discharged from the platform via the plurality of outlet holes 70, 70a, 70b opening into the mate face 22.

[0021] In the embodiment shown in FIG. 3, the platform cooling circuit 30 comprises a serpentine passage formed by a plurality of serially connected flow passes having alternating flow directions. The present example, which is non-limiting, depicts a four-pass serpentine passage comprising flow passes 34, 36, 38, 40. As shown, the flow direction of each of the flow passes 34, 36, 38, 40 may extend generally in an axial direction, i.e., in a forward-to-aft or an aft-to-forward direction, preferably extending for a major portion of the axial length of the platform 6. As a whole, the serpentine passage conducts the coolant K from the inlet 32 in a direction from the pressure side wall 12 toward the pressure side mate face 22. Based on the cooling requirement, the flow passes 34, 36, 38, 40 may have different possible contours, straight or curved. For example, as shown in FIG. 3, the first or upstream- most flow pass 34 may generally follow the contour of the pressure side wall 12 to provide additional cooling to the regions in the platform 6 adjoining the airfoil 10. The last or downstream-most flow pass most proximate to the mate face 22, which is formed by the split channel 40, may extend generally parallel to the mate face 22. Optionally, turbulators 90 may be provided at various locations along the serpentine cooling passage to enhance convective heat transfer between the coolant K and the platform 6. As a further optional feature, one or more of flow passes (in this example, the flow pass 36) may have a variable flow cross-section. In other examples, instead of a serpentine circuit, various other cooling circuit configurations may be implemented.

[0022] An added benefit of the illustrated embodiments is that it allows a targeted cooling of the platform mate face by placing mate face cooling outlet holes as necessary along the mate face. In the shown example, all of the outlet holes 70a, 70b are located at an aft portion of the platform 6, to provide a targeted cooling at the aft region of the mate face 22. To this end, the flow direction of the cooling channel 40 extends in the forward-to-aft direction, with the outlet holes 70a, 70b being located at the downstream end of the respective flow channels 40a, 40b. Furthermore, as shown, at least some of the plurality of outlet holes 70a, 70b may be inclined to the mate face 22 toward an aft direction. The cooling holes may be inclined at an angle to provide additional cooling hole length which will increase the convective cooling of the portion of the blade or vane platform between the cooling channel and the mate face. Having the holes angled also may allow the cooling holes to be sourced as far down the length of the mate face as possible while cooling a more forward portion of the mate face so that the coolant in the cooling channel can pass through as much of the length of the cooling channel as possible thereby reducing the size of the "reduced heat transfer region".

[0023] In an alternate example embodiment, the cooling channel 40 may have a flow direction counter to that illustrated in FIG. 3 (i.e., aft-to-forward), opening into mate face cooling outlet holes located in a forward portion of the platform 6. The number, placement and/or geometry of the mate face cooling outlet holes may be optimized to the specific engine application, being based on several factors including (but not limited to): firing temperature, geometry of the airfoil, pressure, mass flow of the coolant, contouring of the end-wall (platform), etc.

[0024] As described, the illustrated embodiments provide a platform pressure side cooling circuit 30 located entirely on a pressure side section 7 of the platform 6 adjacent to the airfoil pressure side wall 12. Although not shown, the underlying idea may additionally or alternately be applied to a platform suction side cooling circuit located on a suction side section 8 of the platform 6 adjacent to the airfoil suction side wall 14, with outlet holes located on the suction side mate face 24. Turbine components employing the illustrated cooling features may be formed, for example, by an investing casting process using a ceramic casting core. The shown embodiments may also be formed by other manufacturing techniques, for example, by additive manufacturing. Furthermore, embodiments of the present invention may be employed in platforms provided with three-dimensional end-wall contouring. In such an example, as shown in FIG. 6, a cooling channel 40 may be formed so as to follow the contour of the outer surface 6a of the platform, so that a constant thickness "t" is defined between the coolant and the hot gas path. In alternate embodiments, this thickness "t" may be varied, being optimized to specific cooling requirements.

[0025] The illustrated embodiments employ aspects of the present invention on an inner diameter platform of a turbine rotor blade. One skilled in the art would appreciate that the underlying teachings may be applied in a turbine stationary vane at an outer diameter platform and/or at an inner diameter platform.

[0026] 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.