Patent Publication Number: US-10323524-B2

Title: Axial skin core cooling passage for a turbine engine component

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with government support under contract number FA8650-09-D-2923-0021 awarded by the United States Air Force. The government has certain rights in the invention. 
    
    
     TECHNICAL FIELD 
     The present disclosure relates generally to turbine engine components, such as blades, and more specifically to a turbine engine component including an axially aligned skin core cooling passage. 
     BACKGROUND 
     Gas turbine engines, such as those utilized in commercial and military aircraft, utilize a compressor section to draw air into a flow path, a combustor section to mix compressed air with a fuel and ignite the mixture, and a turbine section to expand the resultant combustion products. The expansion of the resultant combustion products drives the turbine section to rotate, which in turn drives the compressor section to rotate. 
     As a result of the exposure to combustion products, components within the turbine section are subject to extreme heat. To prevent heat related fatigue and damage, the turbine components are actively cooled via internal cooling flow paths. Frequently air, or another coolant, is expelled from the internal cooling passages along the surface of the turbine engine component to create a film cooling effect on the exterior surface of the turbine engine component. 
     SUMMARY OF THE INVENTION 
     In one exemplary embodiment a turbine engine component includes a fore edge connected to an aft edge via a first surface and a second surface, a plurality of cooling passages defined within the turbine engine component, and a first skin core passage defined immediately adjacent one of the first surface and the second surface, wherein approximately 100% of coolant entering the first skin core passage is expelled from the turbine engine component at the aft edge. 
     In another exemplary embodiment of the above described turbine engine component, the turbine engine component has an airfoil shaped profile, the fore edge is a leading edge, the aft edge is a trailing edge, the first surface is a suction side and the second surface is a pressure side. 
     In another exemplary embodiment of any of the above described gas turbine components, the first skin core passage extends along the pressure side. 
     In another exemplary embodiment of any of the above described gas turbine components, the first skin core passage defines an axial flow path relative to an axis defined by a turbine engine including the turbine engine component. 
     In another exemplary embodiment of any of the above described gas turbine components, each of the plurality of cooling passages is a section of a single cooling flow path. 
     In another exemplary embodiment of any of the above described gas turbine components, the first skin core passage has a constant axial width. 
     In another exemplary embodiment of any of the above described gas turbine components, the first skin core passage has a varied axial width. 
     In another exemplary embodiment of any of the above described gas turbine components, the axial width of the first skin core passage is minimized at the aft edge. 
     In another exemplary embodiment of any of the above described gas turbine components, the plurality of cooling passages further includes at least one radially aligned skin core passage defining a radial cooling air flow path section relative to an axis defined by the turbine engine including the turbine engine component. 
     In another exemplary embodiment of any of the above described gas turbine components, the radial skin core passage is immediately adjacent one of the first surface and the second surface opposite the one of the first surface and the second surface to which the first skin core passage is immediately adjacent. 
     In another exemplary embodiment of any of the above described gas turbine components, the turbine engine component is one of a blade outer air seal, a combustor liner, a blade and a vane. 
     In another exemplary embodiment of any of the above described gas turbine components, the turbine engine component is a blade in a second or later turbine stage. 
     In one exemplary embodiment a gas turbine engine includes a compressor section, a combustor section fluidly connected to the compressor section by a flowpath, a turbine section fluidly connected to the combustor section by the flowpath, at least one gas turbine engine component exposed to a fluid passing through the flowpath. The at least one gas turbine engine component includes a fore edge connected to an aft edge via a first surface and a second surface, at least one cooling passage defined within the turbine engine component, a first skin core passage defined immediately adjacent the first surface, wherein approximately 100% of coolant entering the first skin core passage is expelled from the turbine engine component at the aft edge. 
     In another exemplary embodiment of the above described gas turbine engine, coolant expelled from the turbine engine component at the aft edge is expelled into the flowpath. 
     In another exemplary embodiment of any of the above described gas turbine engines, each at least one cooling passage and the first skin core passage are sections of a singular cooling flowpath 
     In another exemplary embodiment of any of the above described gas turbine engines, the first skin core passage defines an axial flow relative to an axis defined by the gas turbine engine. 
     In another exemplary embodiment of any of the above described gas turbine engines, the first skin core passage has a constant axial width. 
     An exemplary method for constructing an engine component includes defining a negative image of at least one internal cooling passage and an axial skin core passage connected to the at least one internal cooling passage, casting a material about the negative image, and removing the negative image from a cast component. 
     In a further example of the above exemplary method, defining a negative image includes defining the negative image using a refractory metal core and wherein casting the material about the negative image is an investment casting process. 
     In a further example of any of the above exemplary methods, defining a negative image includes defining the negative image using a ceramic core. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically illustrates an exemplary gas turbine engine. 
         FIG. 2  schematically illustrates a turbine engine component. 
         FIG. 3A  schematically illustrates a cross section of a first example turbine engine component according to  FIG. 2 . 
         FIG. 3B  schematically illustrates a cross section of a second example turbine engine component according to  FIG. 2 . 
         FIG. 4  schematically illustrates a cross section of an alternate example turbine engine component. 
         FIG. 5  schematically illustrates a cross section of another alternate example turbine engine component. 
         FIG. 6  illustrates an exemplary Blade Outer Air Seal including an axially aligned skin core. 
     
    
    
     DETAILED DESCRIPTION OF AN EMBODIMENT 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]^0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
     Multiple components within the turbine section  28  include internal cooling passages for active cooling. Cooling air is typically drawn from the compressor section  24 , such as via a compressor bleed, and provided to the cooled turbine component. Turbine engine components exposed to the hottest temperatures, such as turbine blades and vanes in the first stage aft of the combustor section  26 , are allocated the highest amount of cooling air (referred to as the cooling air budget). Later stages of vanes, blades, blade outer air seals, and other turbine engine components that are further downstream are provided a limited cooling air budget, relative to the cooling air budget of the first stage blade. 
     While film cooling is frequently employed as a cooling method, film cooling produces a significant drop in coolant pressure at the film cooling holes. In a blade, or other turbine engine component, having a limited cooling air budget, the resultant pressure drop can reduce the ability to provide internal cooling downstream of the film cooling holes. 
     With continued reference to  FIG. 1 ,  FIG. 2  illustrates an exemplary turbine second stage blade  100 . The turbine second stage blade  100  includes a blade portion  110  extending from a platform  120  into the primary flow path of the gas turbine engine  20 . A root portion  130  is received in the gas turbine engine static support structure, and maintains the turbine second stage blade  100  in position. The blade portion  110  has a forward edge, referred to as a leading edge  112 , and an aft edge, referred to as a trailing edge  114 . A cross section A-A of the blade portion  110  drawn from the leading edge  112  to the trailing edge  114  has an airfoil shaped profile (illustrated in  FIGS. 3 and 4 ) with a suction surface  116  and a pressure surface  118  connecting the leading edge  112  to the trailing edge  114 . 
     Included within the blade portion  110  are multiple radially aligned internal cooling passages  210  (illustrated in  FIGS. 3 and 4 ). The radially aligned internal cooling passages  210  form an internal cooling air flow path. Along one surface of the blade portion  110 , such as the pressure surface  118 , is an axially aligned skin core passage  220 ,  220   a  (illustrated in  FIGS. 3 and 4 ). The axially aligned skin core passage  220 ,  220   a  defines an axial flow path, and passes cooling air internally along the surface  118 , thereby cooling the surface  118 . Cooling air entering the axially aligned skin core passage  220 ,  220   a  exits the blade portion  110  through slots  119  in the trailing edge  114 . In some example embodiments additional openings  111  connect the axially aligned skin core passage  220 ,  220   a  to a radially outward surface of the tip portion  109 . In such examples, a small portion of the air passing through the axially aligned skin core passage  220 ,  220   a  exits through the openings  111 . The small portion of the air passing through the axially aligned skin core passage  220 ,  220   a  that exits through the openings  111  can be less than 20% of the air in some examples, and less than 5% of the air in some examples. In alternative examples, the slots  119  can be replaced with multiple holes, or a single slot extending the full radial height of the axially aligned skin core passage  220 ,  220   a.    
     In some examples, the blade  100 , illustrated in  FIG. 2 , is created utilizing an investment casting process. In alternative examples, the second stage blade  100  can be created using a process other than investment casting. In either case, the blade  100  is cast around a core defining a negative image of the internal cooling passages of the blade  100 . The core is then removed from the component leaving the cooling passages empty. A skin core passage, such as the above described skin core passage, is a passage formed around a thin sheet of material that conforms with the shape of the surface along which the skin core passage extends. In the example of  FIG. 2 , the blade portion  110  is formed around the thin sheet of material, and the thin sheet is removed from the formed blade portion  110  once the second stage blade  100  has been cast. The result is a thin passage that extends along the surface to which the skin core passage is adjacent. Cooling air passing through the skin core passage  220  absorbs heat from the surface via convection, allowing for the surface to be actively cooled. 
     With continued reference to  FIG. 2 ,  FIG. 3A  illustrates a first example cross section  200  drawn along cross section A-A. In the example cross section  200  of  FIG. 3 , multiple cooling passages  210  are defined between the suction surface  116  and the pressure surface  118 . The cooling passages  210  are radially aligned so that a coolant, such as cooling air, flows through the cooling passage  210  radially relative to an axis defined by the gas turbine engine including the second stage blade  100 . Each of the passages  210  is interconnected with at least one of the other passages  210  such that the passages  210  form a single cooling flow path through the second stage blade  100 . 
     Defined immediately adjacent to the suction surface  116  are multiple radially aligned skin core passages  212 . In alternative examples, a single radially aligned skin core passage may be utilized in place of the multiple passages  212 . In yet further alternate examples, the radially aligned skin core passages  212  are omitted entirely. 
     Immediately adjacent the pressure surface  118  of the second stage blade  100  is an axially aligned skin core passage  220 . The axially aligned skin core passage  220  defines an axial coolant flow path, relative to the gas turbine engine including the second stage blade  100 . Due to the reduced cooling air budget of the second stage blade  100 , relative to a first stage blade, no film cooling holes or film cooling slots are included along the length of the axially aligned skin core passage  220 . The lack of film cooling holes maintains the cooling air pressure throughout the axially aligned skin core passage  220  until the cooling air is expelled at the trailing edge  114  through cooling air outflow slots  119 , or through openings connecting the axially aligned skin core passage  220  to the tip. 
     The axially aligned skin core passage  220  illustrated in  FIG. 3A  extends a majority of the axial length of the pressure surface  118 , prior to being expelled out the slots  119 . With continued reference to  FIG. 3A ,  FIG. 3B  illustrates an alternate example cross section  200  drawn along cross section A-A, where the axially aligned skin core passage  220   a  extends less than 50% of the axial length of the pressure surface  118 . The alternate example configuration illustrated in  FIG. 3B  can be utilized when only the aft portion of the pressure surface  118  requires the increased cooling provided by the skin core passage  220 , or when the cooling budget is insufficient to allow for a full axial length cooling skin core passage  220 . 
     In each of the examples of  FIGS. 3A and 3B , the utilization of the outlet slot  119  at the trailing edge allows approximately 100% of the cooling air entering the axially aligned skin core passage  220 ,  220   a  to be expelled from the second stage blade  100  at the trailing edge. In other words, no cooling air is expelled from the axially aligned skin core passage  220 ,  220   a  upstream of an outlet disposed at an aft most edge of the axially aligned skin core passage  220 ,  220   a.  In alternative examples, the above described openings connecting the tip of the blade to the axially aligned skin core passage  220 ,  220   a  can allow a small portion of the cooling air to be expelled radially outward of the blade at the tip. In some examples, the amount of air expelled at the tip can be less than 20%. In other examples the amount of air expelled at the tip can be less than 5%. 
     With continued reference to  FIGS. 1-3B ,  FIG. 4  illustrates an alternative cross section of the second stage blade  100  constructed using a refractory metal core. As with the example of  FIG. 3A , the blade portion  110  includes a leading edge  112 , a trailing edge  114 , a pressure surface  118  and a suction surface  116 . In the illustrated example of  FIG. 4 , an axially aligned skin core passage  320  extends along the suction surface  116 , and provides cooling airflow as described above with regards to  FIGS. 3A and 3B . The utilization of a refractory metal core for the casting results in an axial skin core passage  320  that has a uniform width  330  for the full length of the axial skin core passage  320 . The uniform width  330  created by the refractory metal core can be extremely small, relative to other methods of creating cooling passages within a turbine engine component. 
     With continued reference to  FIGS. 1-4 ,  FIG. 5  illustrates the cross section of the second stage blade  100  illustrated in  FIG. 4 , cast around a ceramic core rather than a refractory metal core. Internal cooling passages  402  defined by the ceramic core are larger than the internal passages defined by a refractory metal core. Utilization of a ceramic core in the casting process further allows a width  430   a - d  of the skin core passage  420  to be varied along the axial length of the skin core passage  420 . In the example of  FIG. 5 , the width  430   a  at an entryway to the axial skin core passage  420  is maintained constant through a width  430   b  at the midway point through the axial skin core passage  420 . After the midway point the width slightly increases to a maximum width  430   c,  after which the width  430  is decreased to a minimum width  430   d  at the trailing edge  114 . By varying the width  430  of the skin core passage  420 , the speed at which cooling air is passed through the axial skin core passage  420  can be locally accelerated (when the width is narrower) and locally decelerated (when the width is wider) to adjust the convection capabilities of the cooling air passing through the skin core passage  420  for particular hot spots and cold spots. 
     With general reference to  FIGS. 2-5 , an axially aligned skin core passage that passes all coolant entering the skin core passage to an aft most exit to the skin core passage is capable of providing convective cooling to a surface of the turbine engine component without cooling air pressure loss when the cooling air budget provided to the turbine engine component is low. 
     In some examples, an axially aligned skin core passage, such as is described above, can be utilized in conjunction with a cooling flow that has sufficient cooling air budget for film cooling upstream of the axially aligned skin core, or in a turbine engine component that utilizes multiple cooling air flow paths through the turbine engine component where one of the cooling air flow paths has sufficient budget for film cooling, but the cooling air flow path feeding the axially aligned skin core passage lacks sufficient budget. In such an example, it can be beneficial to provide film cooling from the cooling air flow path, or the upstream portion of the cooling air flow path, where there is sufficient cooling air budget. However, inclusion of the axially aligned skin core passage prevents a direct film cooling hole between the higher budget cooling flow and the surface of the turbine engine component receiving film cooling. 
     While described above with regards to blades in general, and a second stage blade in a turbine section of a gas turbine engine in particular, one of skill in the art having the benefit of this disclosure will understand that the above described skin core cooling passage and pedestal arrangement can be applied to any number of actively cooled turbine engine components including, vanes, blades, blade outer air seals, and the like. 
     By way of example,  FIG. 6  illustrates an exemplary blade outer air seal (BOAS)  500 , including an axially aligned skin core passage  510  providing cooling to a radially interior surface  520  of the BOAS  500 . The skin core passage  510  receives cooling air from at least one internal cooling passage  530  that in return receives cooling air from a cooling air inlet  532 . The skin core passage  510  extends axially along the radially inward surface  520  of the BOAS  500 , and defines an axial flowpath. An outlet of the skin core passage  510  is positioned at an aft edge  550  of the BOAS  500 . The outlet expels cooling air from the skin core passage  510  into a primary flowpath. The positioning of the cooling air inlet  532 , and the internal cooling air passages  530  may be varied depending on the particular needs and configuration of a given engine. 
     It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.