Abstract:
A ceramic casting core, including: a plurality of rows ( 162, 166, 168 ) of gaps ( 164 ), each gap ( 164 ) defining an airfoil shape; interstitial core material ( 172 ) that defines and separates adjacent gaps ( 164 ) in each row ( 162, 166, 168 ); and connecting core material ( 178 ) that connects adjacent rows ( 170, 174, 176 ) of interstitial core material ( 172 ). Ends of interstitial core material ( 172 ) in one row ( 170, 174, 176 ) align with ends of interstitial core material ( 172 ) in an adjacent row ( 170, 174, 176 ) to form a plurality of continuous and serpentine shaped structures each including interstitial core material ( 172 ) from at least two adjacent rows ( 170, 174, 176 ) and connecting core material ( 178 ).

Description:
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT 
     Development for this invention was supported in part by contract Award Number DE-SC0001359 awarded by the United States Department of Energy Office of Science (SBIR) to Mikro Systems, Inc. of Charlottesville, Virginia. Accordingly, the United States Government may have certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     The invention relates to a casting core for forming cooling channels in a gas turbine engine component. In particular the invention relates to a casting core for forming serpentine cooling channels defined by rows of aerodynamic structures. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engines create combustion gas which is expanded through a turbine to generate power. The combustion gas is often heated to a temperature which exceeds the capability of the substrates used to form many of the components in the turbine. To address this, the substrates are often coated with thermal barrier coatings (TBC) and also often include cooling passages throughout the component. A cooling fluid such as compressed air created by the gas turbine engine&#39;s compressor is typically directed into an internal passage of the substrate. From there, it flows into the cooling passages and exits through an opening in the surface of the component and into the flow of combustion gas. 
     Certain turbine components are particularly challenging to cool, such as those components having thin sections. The thin sections have relatively large surface area that is exposed to the combustion gas, but a small volume with which to form cooling channels to remove the heat imparted by the combustion gas. Examples of components with a thin section are those having an airfoil, such as turbine blades and stationary vanes. The airfoil usually has a thin trailing edge. 
     Various cooling schemes have been attempted to strike a balance between the competing factors. For example, some blades use structures in the trailing edge, where cooling air flowing between the structures in a first row is accelerated and impinges on structures in a second row. A faster flow of cooling fluid will more efficiently cool than will a slower flow of the same cooling fluid. This may be repeated to achieve double impingement cooling, and repeated again to achieve triple impingement cooling, after which the cooling air may exit the substrate through an opening in the trailing edge, where the cooling air enters the flow of combustion gas passing thereby. The impingement not only cools the interior surface of the component, but it also helps regulate the flow. In particular it may create an increased resistance to flow along the cooling channel and this may prevent use of excess cooling air. 
     For cost efficient cooling design the trailing edge is typically cast integrally with the entire blade using a ceramic core. The features and size of the ceramic core are important factors in the trailing edge design. A larger size of a core feature makes casting easier, but the larger features are not optimal for metering the flow through the crossover holes to achieve efficient cooling. In the trailing edge, for example, since cavities in the substrate correspond to core material, a crossover holes between the adjacent pin fins in a row corresponds to sparse casting core material in that location of the casting. This, in turn, leads to fragile castings that may not survive normal handling. To achieve acceptable core strength the crossover holes must exceed a size optimal for cooling efficiency purposes. However, the crossover holes result in more cooling flow which is not desirable for turbine efficiency. Consequently, there remains room in the art for improvement. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in the following description in view of the drawings that show: 
         FIG. 1  is a cross sectional side view of a prior art turbine blade. 
         FIG. 2  shows a core used to manufacture the prior art turbine blade shown in  FIG. 1 . 
         FIG. 3  is a cross sectional end view of a turbine blade. 
         FIG. 4  is a partial cross sectional side view along  4 - 4  of the turbine blade of  FIG. 3  showing the cooling channels disclosed herein. 
         FIG. 5  is a close up view of the cooling arrangement of  FIG. 4 . 
         FIG. 6  shows a portion of a core used to manufacture the turbine blade of  FIG. 4 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present inventors have devised an innovative cooling arrangement for use in a cooled component and a casting core that may be used to effect the cooling arrangement when a casting process is used to create the component. The component may alternately be manufactured via machining, or using sheet material. Sheet material may be particularly useful in a component such as a transition duct. The cooling arrangement may include cooling channels characterized by a serpentine or zigzag flow axis, where the cooling channel walls are defined by rows of discrete aerodynamic structures that form continuous cooling channels having discontinuous walls. The aerodynamic structures may be airfoils or the like. The cooling channels may further include other cooling features such as turbulators, and may further be defined by other structures such as pin fins or mesh cooling passages. The cooled component may include items such as blades, vanes, and transition ducts etc that have thin regions with relatively larger surface area. An example of such a thin area is a trailing edge of the blade or vane, but is not limited to these thin areas or to these components. 
     The cooling arrangement disclosed herein enables highly efficient cooling by providing increased surface area for cooling and sufficient resistance to the flow of cooling air while also enabling a core design of greater strength. Traditional flow restricting impingement structures regulated an amount of cooling fluid used by restricting the flow, and this restriction also accelerated the flow in places. A faster moving flow provides a higher heat transfer coefficient, which, in turn, improves cooling efficiency. In the cooling arrangement disclosed herein, the serpentine cooling channels provide sufficient resistance to the flow to obviate the need for the flow restricting effect of the traditional impingement structures. The increased surface area and associated increase in cooling channel length yields an increase in cooling, despite the relatively slower moving cooling fluid having a relatively lower heat transfer coefficient when compared to the faster moving fluid of the impingement-based cooling schemes. The result is that the cooling arrangement disclosed herein yields an increase in overall heat transfer because the positive effect of the increase in surface area more than overcomes the negative effect of the decreased heat transfer coefficient. The satisfactory flow resistance offered by the serpentine shape of the cooling channel is sufficient to regulate the flow and thereby enable the cooling arrangement, with or without the assistance of an array of pin fins or the like. Experimental data indicated upwards of a 40 degree Kelvin temperature drop at a point on the surface of the blade when the cooling arrangement disclosed herein is implemented. 
       FIG. 1  shows a cross section of a prior art turbine blade  10  with an airfoil  12 , a leading edge  14  and a trailing edge  16 . The prior art turbine blade  10  includes a trailing edge radial cavity  18 . Cooling fluid  20  enters the trailing edge radial cavity  18  through an opening  22  in a base  24  of the prior art turbine blade  10 . The cooling fluid  20  travels radially outward and then travels toward exits  26  in the trailing edge  16 . As the cooling fluid  20  travels toward the trailing edge exit  26  it encounters a first row  28  and a second row  30  of crossover hole structures  32 . The cooling fluid  20  flows through relatively narrow crossover holes  34  between the crossover hole structures  32  of the first row  28 , which accelerates the cooling fluid which, in turn, increases the heat transfer coefficient in a region where the accelerated fluid flows. The cooling fluid  20  impinges on the crossover hole structures  32  of the second row  30 , and is again accelerated through crossover holes  34  between the crossover hole structures  32  of the second row  30 . Here again the accelerated fluid results in a higher heat transfer coefficient in the region of accelerated fluid flow. The cooling fluid  20  then impinges on a final structure  36  which keep the fluid flowing at a fast rate before exiting the prior art turbine blade  10  through the trailing edge exits  26  where the cooling fluid  20  joins a flow of combustion gas  38  flowing thereby. Between the trailing edge radial cavity  18  and the trailing edge exit  26  individual flows between the crossover hole structures  32  may be subsequently split when impinging another crossover hole structures  32  or final structure  36 , and split flows may be joined with other adjacent split flows. Consequently, it is difficult to describe the cooling arrangement in the prior art trailing edge  16  as continuous cooling channels; it is better characterized as a field of structures that define discontinuous pathways where individual flows of cooling fluid  20  split and merge at various locations throughout. 
       FIG. 2  shows a prior art core  50  with a core leading edge  52  and a core trailing edge  54  and a core base  55 . During manufacture a substrate material (not shown) may be cast around the prior art core  50 . The solidified cast material becomes the substrate of the component. The prior art core  50  is removed by any of several methods known to those of ordinary skill in the art. What remains once the prior art core  50  is removed is a hollow interior that forms the trailing edge radial cavity  18  and the crossover holes  34 , among others. For example, core crossover hole structure gaps  56  are openings in the prior art core  50  which will be filled with substrate material and form crossover hole structures  32  in the prior art blade  10  (or vane etc). Conversely, core crossover hole structures  58  between the core crossover hole structure gaps  56  will block material in the substrate so that once the prior art core  50  is removed the crossover holes  34  will be formed. It can be seen that the core crossover hole structures  58  are relatively small in terms of depth (into the page) and height (y axis on the page) and provide a weak regions  60 ,  62 ,  64  that correspond to locations in the prior art core  50  that form the first row  28 , the second row  30 , and the row of final structures  36  in the finished prior art turbine blade  10 . These weak regions  60 ,  62 , and  64  may break prior to casting of the substrate material and this is costly in terms of material and lost labor etc. 
       FIG. 3  is a cross sectional end view of a turbine blade  80  having the cooling arrangement  82  disclosed herein in a trailing edge  84  of the turbine blade  80 . The cooling arrangement  82  is not limited to a trailing edge  84  of a turbine blade  80 , but can be disposed in any location where there exists a relatively large surface area to be cooled. In the exemplary embodiment shown the cooling arrangement  82  spans from the trailing edge radial cavity  86  to the trailing edge exits  88 . 
       FIG. 4  is a partial cross sectional side view along  4 - 4  of the turbine blade  80  of  FIG. 3  showing cooling channels  90  of the cooling arrangement  82 . In the exemplary embodiment shown the cooling channels  90  are defined by a first row  92 , a second row  94 , and a third row  96  of flow defining structures  98  and are continuous and discrete paths for a cooling fluid. However, each cooling channel  90  is not continuously bounded by flow defining structures  98 . Instead, between rows  92 ,  94 ,  96  of flow defining structures  98  each cooling channel  90  is free to communicate with an adjacent cooling channel  90 . Downstream of the cooling channels  90  there may be an array  100  of pin fins  102  or other similar structures used to enhance cooling, meter the flow of cooling fluid, and provide strength to both the turbine blade  80  and the prior art core  50 . In the exemplary embodiment shown the flow defining segments  98  take the form of an airfoil, but other shapes may be used. 
       FIG. 5  is a close up view of the cooling arrangement  82  of  FIG. 4 . Each cooling channel  90  includes at least two segments where the cooling channel is bounded by flow defining structures  98  that provide bounding walls. In between segments the cooling channel  90  may be unbounded by walls where cross paths  104  permit fluid communication between adjacent cooling channels  90  and contribute to an increase in surface area available for cooling inside the turbine blade  80 . The cooling channels may open into the array  100  of pin fins  102 . In the exemplary embodiment shown there are three rows  92 ,  94 ,  96 , of flow defining structures  98 , and hence three segments per cooling channel  90 . 
     The first row  92  of flow defining structures  98  defines a first segment  110  having a first segment inlet  112  and a first segment outlet  114 . In the first row  92  a first wall  116  of the cooling channel  90  is defined by a suction side  118  of the flow defining structure  98 . A second wall  120  of the cooling channel  90  is defined by a pressure side  122  of the flow defining structure  98 . Between the first row  92  and the second row  94  the cooling channel is not bounded by walls, but is instead open to adjacent channels via the cross paths  104 . 
     The second row  94  of flow defining structures  98  defines a second segment  130  having a second segment inlet  132  and a second segment outlet  134 . In the second row  94  the first wall  116  of the cooling channel  90  is now defined by a pressure side  122  of the flow defining structure  98 . The second wall  120  of the cooling channel  90  is now defined by the suction side  118  of the flow defining structure  98 . Between the second row  94  and the third row  96  the cooling channel is not bounded by walls, but is instead open to adjacent channels via the cross paths  104 . 
     The third row  96  of flow defining structures  98  defines a third segment  140  having a third segment inlet  142  and a third segment outlet  144 . In the third row  96  the first wall  116  of the cooling channel  90  is defined by a suction side  118  of the flow defining structure  98 . The second wall  120  of the cooling channel  90  is defined by a pressure side  122  of the flow defining structure  98 . The cooling channel  90  ends at the third segment outlet  144 , where the cooling channel may open to the array  100  of pin fins  102 . The array  100  of pin fins  102  may or may not be included in the cooling arrangement  82 . 
     Unlike conventional impingement based cooling arrangements, the instant cooling arrangement  82  aligns the outlets and inlets of the segments so that cooling air exiting an outlet is aimed toward the next segment&#39;s inlet. This aiming may be done along a line of sight (mechanical alignment), or it may be configured to take into account the aerodynamic effects present during operation. In a line of sight/mechanical alignment an axial extension  152  of an outlet in a flow direction will align with an inlet of the next/downstream inlet. An aerodynamic alignment may be accomplished, for instance, via fluid modeling etc. In such instances an axial extension of an outlet may not align exactly mechanically with an inlet of the next/downstream inlet, but in operation the fluid exiting the outlet will be directed toward the next inlet in a manner that accounts for aerodynamic influences, such as those generated by adjacent flows, or rotation of the blade etc. It is understood that the cooling fluid may not exactly adhere to the path an axial extension may take, or a path on which it is aimed in an aerodynamic alignment, but it is intended that the fluid will flow substantially from an outlet to the next inlet. Essentially, the fluid may be guided to avoid or minimize impingement, contrary to the prior art. 
     This aiming technique may also be applied to cooling fluid exiting the third segment outlet  144  at the end of the cooling channel  90 . In particular an axial extension of the third segment outlet  144  may be aimed between pin fins  102  in a first row  146  of pin fins  102  in the array  100 . Likewise the flow exiting the third segment outlet  144  may be aerodynamically aimed between the pin fins  102  in the first row  146 . Still further, downstream rows of pin fins may or may not align to permit an axial extension of the third segment outlet  144  to extend uninterrupted all the way through the trailing edge exits  88 . The described configuration results in a cooling channel  90  with a serpentine flow axis  150 . The serpentine shape may include a zigzag shape. 
     The cooling channels  90  may have turbulators to enhance heat transfer. In the exemplary embodiment shown the cooling channels  90  include mini ribs, bumps or dimples  148 . Alternatives include other shapes known to those of ordinary skill in the art. These turbulators increase surface area and introduce turbulence into the flow, which improves heat transfer. 
       FIG. 6  shows an improved portion  160  of an improved core, the improved portion  160  being for the trailing edge radial cavity  86  and designed to create the cooling arrangement  82  disclosed herein. (The remainder of the improved core would remain the same as shown in  FIG. 2 .) A first row  162  of core flow defining structure gaps  164 , a second row  166  of core flow defining gaps  164 , and a third row  168  of core flow defining gaps  164  are present in the improved core portion  160  where the first row  92 , the second row  94 , and the third row  96  of flow defining structures  98  respectively will be formed in the cast component. A first row  170  of interstitial core material  172  separates the core flow defining structure gaps  164  in the first row  162  from each other. A second row  174  of interstitial core material  172  separates the core flow defining structure gaps  164  in the second row  166  from each other. A third row  176  of interstitial core material  172  separates the core flow defining structure gaps  164  in the third row  166  from each other. Each row ( 170 ,  174 ,  176 ) of interstitial core material is connected to an adjacent row with connecting core material  178  that spans the rows ( 170 ,  174 ,  176 ) of interstitial core material. A first row  180  of core pin fin gaps  182  begins an array  184  of pin fin gaps  182  where the first row  146  of pin fins  102  and the array  100  of pin fins  102  will be formed in the cast component. Also visible are core turbulator features  188  where mini ribs, bumps or dimples  148  will be present on the cast component. The improved portion  160  may also include surplus core material  186  as necessary to aid the casting process. 
     When compared to the trailing edge portion of the prior art core  50  of  FIG. 2 , it can be seen that the improved core portion  160  is structurally more sound than the trailing edge portion of the prior art core  50 . In particular, the improved core portion  160  does not have the weak regions  60 ,  62 ,  64  which include material that is relatively small in terms of depth (into the page) and height (y axis on the page). Instead, the rows  170 ,  174 ,  176  of interstitial core material  172  are present between the core flow defining structure gaps  162  in the improved core portion, and the interstitial core material  172  has a same depth as the flow defining structure gaps  162  themselves (i.e. the interstitial core material  172  is as thick as the bulk of the improved core portion  160 ) and thus the improved core portion  160  is stronger than the prior art design. 
     Stated another way, a first region  190  immediately upstream of a respective row of the interstitial core material  172  has a first region thickness. A second region  192  immediately downstream of a respective row of the interstitial core material  172  has a second region thickness. The interstitial core material  172  between the first region and the second region has an upstream interstitial core material thickness that matches the first region thickness because they blend together at an upstream end of the interstitial core material  172 . The interstitial core material  172  has a downstream interstitial core material thickness that matches the second region thickness because they blend together at a downstream end of the interstitial core material  172 . The interstitial core material  172  maintains a maximum thickness between the upstream end and the downstream end. This configuration is the same for all of the rows  170 ,  174 ,  176  of interstitial core material  172 . Since there is no reduction in thickness of the improved core portion  160  where the interstitial core material  172  is present, the improved core portion  160  is much stronger than the prior art core portion  50 . This reduces the chance of core fracture and provides lower manufacturing costs associated there with. Furthermore, the relatively larger cooling passages disclosed herein are less susceptible to clogging from debris that may find its way into the cooling passage than the crossover holes of the prior art configuration. 
     The cooling arrangement disclosed herein replaces the impingement cooling arrangements of the prior art which accelerate the flow to increase the cooling efficiency with a cooling arrangement having serpentine cooling channels. The serpentine channels provide sufficient resistance to flow to enable efficient use of compressed air as a cooling fluid, and the increased surface area improves an overall heat transfer quotient of the cooling arrangement. Further, the improved structure can be cast using the casting core with improved core strength. As a result, cooling efficiency is improved and manufacturing costs are reduced. Consequently, this cooling arrangement represents improvements in the art. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.