Patent Application
20100322767
The present invention relates generally to techniques for reducing cracks in platforms of gas turbine blades and more specifically to a turbine blade having a plurality of cooling holes formed in the platform.
The turbine section of gas turbine engines typically comprises multiple sets or stages of stationary airfoils, known as nozzles or vanes, and moving airfoils, known as rotor blades or buckets.
In one such solution, an undercut is machined into the blade platform along the pressure side of the blade. This proposed solution purports to reduce the total stress level in this region of high stress. This approach has been implemented on both turbine and compressor blades as both a field repair and a design modification. If a stress reduction is achieved in the platform region, the concern is whether the undercut results in a high stress within the grooved region where material is removed. In other words, the success of the strategy depends on whether a stress reduction in an existing high-stress region can be achieved without creating a new area of high stress within the blade.
There are two primary concerns raised with platform undercuts. First, whether the undercut will be effective in reducing the stress in the platform. Second, whether the stress concentration occurring in the undercut will be so high that it offsets the benefit of the undercut to the platform region. Undercut solutions have had difficulty striking a balance between these two concerns. It is desired to have a solution which effectively reduces the stress in the platform and, thereby, the potential for formation of cracks, TBC delamination, and, in the worst case, breakage and separation of significant portions of the platform altogether, and which does not add additional stresses to the blade. The present invention seeks to solve this problem.
In one embodiment, the present invention is directed to a turbine blade and limits platform cracking. The turbine blade has of an airfoil connected to a platform in the blade root region. The airfoil and the platform share a common cooling passage, which may include one or more cooling channels or paths. The turbine blade configuration limits platform damage, including but not limited to cracking, removal of the TBC layer, and breakage and loss of blade material, through the influence of at least one cooling hole which extends from an outside edge of the platform, through the platform and to the common cooling passage. In one embodiment, multiple cooling holes are formed in the platform at an angle to the outside edge of the platform which is approximately 90° and in another embodiment at approximately 45°. In one embodiment, the common cooling passage has one or more serpentine cooling circuits, and each of the cooling holes extends to a distinct channel within the serpentine cooling circuit. In another embodiment, the common cooling passage includes a plurality of generally parallel cooling veins extending through the platform and airfoil to the airfoil tip, which may be formed by a Shaped-Tube Electolytic Machining (STEM) drilling process for example. In this case, each of the platform cooling holes extends through the platform from the platform edge to a distinct parallel cooling vein in the blade. In one embodiment, the platform cooling holes are formed at the approximate midpoint of the defined thickness of the platform. The cooling holes may be generally cylindrical in shape with a diameter approximately 50% of the platform thickness.
In another embodiment, the present invention is directed to a method of limiting damage to the platform of a turbine blade having an airfoil connected to the platform in a blade root region. The method includes the step of forming at least one cooling hole in the platform which extends from an outside edge of the platform to a cooling passage within the platform which passes through and is shared with the airfoil. In one embodiment, multiple cooling holes are formed in the platform. In a more specific embodiment, there may be four cooling holes formed in the platform. The cooling holes can be formed at the angles, having the location and geometric dimensions as described in the immediately preceding paragraph. The cooling holes can be formed by a number of known processes, but in at least one embodiment is formed by an electro-discharge machining (EDM) process.
The present invention has application in both the manufacture of a new turbine blade as well as in the repair of an existing turbine blade not having cooling holes formed in the platform region thereof. In the latter case, the one or more cooling holes can be formed without having to remove any existing TBC layer that may have been applied on the blade.
The following drawings form part of the present specification and are included to further demonstrate certain aspects of the present invention. The present invention may be better understood by reference to one or more of these drawings in combination with the description of embodiments presented herein. However, the present invention is not intended to be limited by the drawings.
The present invention will now be described with reference to the following exemplary embodiments. Referring now to
The airfoil 206 is defined by a concave side wall 210, a convex side wall 208, a leading edge 212 and an opposite trailing edge 214; the leading and trailing edges being the two areas where the concave side wall and convex side wall meet. The airfoil 206 has a root 216 which is proximate the platform 204 and a tip (or shroud) 218 which is distal from the platform. As with prior art turbine blades, air is supplied to the inside cavity (not shown) of the airfoil 206 from the compressor to cool the airfoil. The cooling air may exit through a plurality of cooling holes 220, at least some of which may be formed in the trailing edge 214.
In accordance with the present invention, the platform 204 has a plurality of cooling holes 230 formed therein on the pressure (concave) side of the airfoil 206, which is the region of the platform that is susceptible to high stresses and often cracks, including delaminating of coating, when present, and separation or breakage of blade base material in extreme cases. In one embodiment, four such cooling holes 230 are formed in the platform 204. The platform cooling holes 230 may be formed by an EDM process. Alternatively, the platform cooling holes 230 can be formed via STEM process or electro-chemical (ECM) drilling process or other similar machining process. The process utilized to form the cooling holes 230 may be selected to avoid removal of the TBC layer formed on the turbine blade. In one embodiment, the platform cooling holes 230 are generally cylindrical in shape, with center axes generally parallel to the lower surface and the upper surface. The cross-section of a platform cooling hole 230 at an outside edge of the platform 204 may span approximately 50% of the platform thickness, or the platform cooling holes 230 may have a diameter of approximately 50% of the thickness of the platform 204. The platform cooling holes 230 may also be formed at the approximate mid-point of the thickness of the platform 204, i.e., the centers of the cross-section of the platform cooling holes 230 at the outside edge of the platform 204 are aligned at the mid-point of the platform thickness so that an equal amount of platform material is left above and below the platform cooling holes 230.
In one embodiment, the platform cooling holes 230 are formed at an angle to the outside edge of the platform 204 into which they are formed, which is best seen in
In another embodiment (shown in
Without limiting the invention to a particular theory or mechanism of action, it is nevertheless currently believed that the overall cooling flow may increase and the internal cooling flow may be re-distributed as a consequence of adding the platform cooling holes 230. Table I lists the cooling mass flow which may occur as a result of adding the platform cooling holes 230 to an example first stage turbine blade with serpentine cooling passages.
As shown in the table, the cooling flow in the leading serpentine cooling circuit may be ˜0.7% more than the prior art blade configuration, and the cooling flow in the trailing serpentine cooling circuit may increase by ˜1.2%. The total cooling flow may increase only marginally (by ˜0.9%) with the drilling of four platform cooling holes 230. The cooling flow of the leading three platform cooling holes 230 may be 6.1, 5.8, and 6.5 pound mass per hour (lbm/hr), respectively. For the 4th platform-cooling hole 230, which branches from the trailing serpentine passage, the flow rate may be 6.1 lbm/hr. The total platform cooling flow may be 24.4 lbm/hr, or about 2.5% of total cooling flow available to the bucket.
The resulting surface temperature distributions of the modified blade with platform cooling holes 230, according to one embodiment of the invention, and a prior art blade are shown in
Further examining these results indicates that the benefit of the proposed platform cooling strategy is likely to be at least twofold. Through the additional convective cooling and conduction, the gross reduction of the temperature in the platform region should favorably lower the temperature gradients near the juncture of platform and trailing edge lowermost cooling hole, which is particularly susceptible to cracking, as indicated in
Equivalent and axial stress distributions of the blade modified with platform cooling, according to one embodiment of the invention, are plotted in
In summary, the platform cooling hole modifications of the present invention may be effective in both reducing the temperatures and stresses in the cooled platform region. Moreover, they may provide additional benefits in lowering the thermal gradient near the juncture of platform and trailing edge, and consequentially reduce the stress at the trailing edge lowermost cooling hole. Based on a comparison to the results of the baseline analysis, it is therefore considered as a viable design modification—to be utilized in the course of forming a new turbine blade—and/or recommended to implement during repair and refurbishment of blades.
The terms “holes,” “passages,” “veins,” “channels,” and the like are each used to describe conduits for the flow of air or other cooling fluid. The use of different words for the various conduits is not intended to be limiting in any way, but instead is to assist the reader in fully understanding the interrelation between the various conduits.
Therefore, the present invention is well adapted to attain the ends and advantages mentioned as well as those that are inherent therein. The particular embodiments disclosed above are illustrative only, as the present invention may be modified and practiced in different but equivalent manners apparent to those skilled in the art, having the benefit of the teachings herein. For example, as those of ordinary skill in the art will appreciate a different number of platform cooling holes 230 may be implemented, such platform cooling holes 230 may be formed at a different angle than that disclosed herein, and such platform cooling holes 230 may be oriented at a different location within the thickness of the platform. Furthermore, no limitations are intended to the details of construction or design herein shown, other than as described in the claims below. It is therefore evident that the particular illustrative embodiments disclosed above may be altered or modified and all such variations are considered within the scope and spirit of the present invention. Also, the terms in the claims have their plain, ordinary meaning unless otherwise explicitly and clearly defined by the patentee.
