Abstract:
A component for a gas turbine engine has an airfoil with internal cooling channels for delivering air from a radially outer end of the airfoil toward a radially inner end of the airfoil. The cooling channels are separated from adjacent cooling channels by sets of at least two disconnected wall segments.

Description:
[0001]    This invention was made with government support under Contract No. F33615-03-D-2354 (DO:0009) awarded by the United States Air Force. The government therefore has certain rights in this invention. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    This application relates to segmented cooling cavities for use within a turbine component having an air cooled blade outer air seal. 
         [0003]    Gas turbine engines are known, and typically include a compressor compressing air and delivering it downstream into a combustion section. The air is mixed with fuel in the combustion section and combusted. Products of this combustion pass downstream over turbine rotors, driving the turbine rotors to rotate and create power. 
         [0004]    Typically, the turbine rotors include a plurality of removable blades. Blade outer air seals are positioned adjacent a radially outer portion of the blades, to provide a tight clearance across the rotors such that air is restricted to flow over the rotors, rather than bypassing them. These blade outer air seals are subjected to a very harsh and hot environment due to the products of combustion. Thus, it is known to deliver cooling air to blade outer air seals. 
         [0005]    A blade outer air seal typically has a leading edge and a trailing edge, defined in terms of the turbine gas flow direction. The pressure across the blade outer air seal drops since the products of combustion are transferring their pressure into energy to drive the turbine rotors. Thus, at the leading edge, the pressure is significantly higher than at the trailing edge. As an example, the pressure could be as much as twice as high as the leading edge as it is at the trailing edge. 
         [0006]    This causes some issues with regard to the flow of cooling air within the blade outer air seal. In some prior art blade outer air seals, a single large cooling channel extends across the entire axial length of the seal, from leading to trailing edge. In such an arrangement, less air will exit at the leading edge compared to the trailing edge. This is because the cooling air will seek the lowest pressure, and will thus tend to flow more toward the trailing edge. 
         [0007]    One solution to this problem has been the use of a plurality of separate cooling channels spaced along the length of the airfoil. These separate channels do limit the effect of the pressure differential between leading edge and the trailing edge. However, the use of the separate channels complicates the manufacture of the airfoil. 
         [0008]    Most components for gas turbine engines containing an airfoil are formed by a loss core molding process. In such a process, a core is initially formed of a particular material. That core is inserted into a mold, and molten metal is directed into the mold and around the core. After the molten metal has hardened, the material of the core is leached away, leaving a cavity where the core once sat. This is typically the manner in which cooling channels are formed. To form the separate channels as mentioned above, the core must have a plurality of separate core members, or alternatively, a plurality of spaced core fingers. The use of several cores is complex, and the use of a single core with separate spaced fingers is not as structurally robust as may be desirable. 
       SUMMARY OF THE INVENTION 
       [0009]    In the disclosed embodiment of this invention, a plurality of segmented cooling channels are formed within a blade outer air seal for a gas turbine engine by disconnected wall members. The wall members have overlapping extents, but do not directly contact each other. In this manner, a single robust core member can be utilized to form multiple semi-discrete channels for one or all cooling cavities within a blade outer air seal. 
         [0010]    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 
         [0011]      FIG. 1  schematically shows a gas turbine engine. 
           [0012]      FIG. 2A  schematically shows one portion of a prior art blade outer air seal. 
           [0013]      FIG. 2B  is a second view of  FIG. 2A . 
           [0014]      FIG. 2C  shows a problem with a second prior art blade outer air seal. 
           [0015]      FIG. 3  shows a cross-sectional view of an inventive blade outer air seal. 
           [0016]      FIG. 4  is a first view of a mold for forming the inventive blade outer air seal. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0017]    A gas turbine engine  10 , such as a turbofan gas turbine engine, circumferentially disposed about an engine centerline  12 , is shown in  FIG. 1 . The engine  10  includes a fan  14 , compressors  15  and  16 , a combustion section  18  and a turbine section  20 . Turbine section  20  includes rotors  13  and  15 . As is well known in the art, air compressed in the compressors  15  and  16  is mixed with fuel and burned in the combustion section  18 , and expanded across turbine rotors  13  and  15 . Turbine rotors  13  and  15  rotate in response to the expansion, driving the compressors  15  and  16 , and fan  14 . Turbine rotor  13  and  15  comprise alternating rows of rotary airfoils or blades  24  spaced from static airfoils or vanes  26 . This structure is shown quite schematically in  FIG. 1 . While one example gas turbine engine is illustrated, it should be understood this invention extends to any other type gas turbine engine for any application. 
         [0018]      FIGS. 2A and 2B  show the circumferential edge portion of an existing blade outer air seal  50 . As shown, a plurality of cooling channels  52  are separated by spaced walls  54 . To form the separate channels  52 , a plurality of separate cores must be utilized during a molding process. The channels may be circumferentially or radially oriented, depending on the details of the particular cooling scheme. In the figures shown, the channels are oriented circumferentially with air exiting in the gap between seal segments. 
         [0019]    On the other hand, if the separate cooling channels  52  are not used then a problem as illustrated in  FIG. 2C  will occur. As shown, a core for a blade outer air seal  51  has a single large cooling channel  55 , and a leading edge  56  and a trailing edge  58 . The pressure drop between the leading edge  56  and the trailing edge  58  may be dramatic. It is not unusual for pressures to be approximately half adjacent the trailing edge as they are at the leading edge. Thus, as shown schematically in  FIG. 2C , a good deal of the cooling air directed into the blade outer air seal  51  will tend to flow toward the trailing edge  58 , and smaller quantities of air reach the leading edge  56 . On the other hand, the cooling air may be more important adjacent the leading edge  56 , if the leading edge is hotter than the trailing edge. Adjusting exit hole sizes may aid in adjusting the flow somewhat, but this provides only limited control and may produce other problems such as plugging of tiny holes. 
         [0020]    As mentioned above, one known solution to address this problem is the use of the separate cooling channels (see blade outer seal  50  in  FIG. 2B ); however, they are somewhat complex to form. 
         [0021]      FIG. 3  schematically shows an inventive blade outer air seal  60  having a unique cooling scheme. The walls between separate cooling channels  61  are formed by spaced cooling wall segments  62  and  65 . The upstream cooling wall segments  62  have an inner end  64  which overlaps with an outer end  66  of a downstream wall segment  65 . Note that in this case, the terms upstream and downstream refer to the flow of the cooling air inside the seal, not the gas flow in the turbine which define the leading and trailing edge of the part. A gap  67  between the wall segments  62  and  66  will allow a single core to be used to form the blade outer air seal  60 , as will be explained below. Now, the benefits of the multi-channel cooling scheme as mentioned above are achieved in that the separating wall segments  62  and  65  do not allow the air to move too far downstream. Gap  67  presents a small and torturous path, so it is unlikely any significant volume of air would move back through the gap  67 . Air would not move from the trailing edge channel downstream toward the adjacent leading edge channel due to the pressure differential. 
         [0022]      FIG. 4  shows a mold system  70  for forming the blade outer seal  60 . Of course, the mold system  70  is shown extremely schematically. A mold housing  72  has an inlet  71  for receiving a molten metal. A core  74  is formed having a plurality of slots  76  and  78 . As a worker in the art of forming gas turbine engine air foil components would understand, the core  74  exists in areas that will provide a space in the final blade outer air seal. The slots  76  and  78  will form the wall segments  62  and  65 . Metal is injected through the inlet  71  into the mold  72 , and is deposited around the core  74 . The core is later leached away, leaving an internal flow structure such as shown in  FIG. 3 . 
         [0023]    While the channels are shown as circumferential channels at an edge periphery of a blade outer air seal, these concepts can also be used in radially extending channels, and also in components other than blade outer air seals. 
         [0024]    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.