Abstract:
A blade outer air seal is provided with a plurality of distinct cooling circuit schemes. Preferably, compact heat exchanger structures are utilized, and can be individually tailored to the particular location along the blade outer air seal. As an example, a greater pressure ratio exists between the products of combustion and the cooling air at the trailing edge than would be found at the leading edge. The present invention takes advantage of this distinction by utilizing cooling schemes that have a greater pressure drop at the trailing edge than the cooling schemes utilized closer to the leading edge.

Description:
This invention was made with government support under Contract No. F33615-03-D-2354-0002 awarded by the United States Air Force. The government therefore has certain rights in this invention. 
    
    
     BACKGROUND OF THE INVENTION 
     This application relates to an improved cooling circuit for a blade outer air seal, in which a plurality of distinct cooling schemes are utilized. 
     Gas turbine engines are provided with a number of functional sections, including a fan section, a compressor section, a combustion section, and a turbine section. Air and fuel are combusted in the combustion section. The products of the combustion move downstream, and pass over a series of turbine rotors, driving the rotors to create power. 
     It is desirable to have the bulk of the products of combustion pass over the turbine blade. Thus, a seal is placed circumferentially about the turbine rotors slightly radially spaced from a radially outer surface of the turbine blades. The seal is in a harsh environment, and must be able to withstand high temperatures. To address the high temperatures, the seal is typically provided with internal cooling channels. Air circulates through the cooling channels to cool the seal. 
     In the prior art, one type of cooling scheme has been utilized across the seal. However, the cooling challenges faced across the seal vary. As an example, the seal extends from a leading edge to a trailing edge. A pressure ratio between the cooling air and the working air is low at the leading edge, and greater at the trailing edge. Even so, the prior art has not tailored the cooling channels to the location. Further, the prior art has typically used only relatively large cooling channels in the blade outer air seals. 
     More recently, compact heat exchanger cooling schemes (or microcircuit cooling channels) have been developed, which utilize relatively thin and small passages to convey cooling air through a body. These compact heat exchangers are formed by lost core molding techniques. While these techniques provide efficient and effective cooling, they have not been applied to a cool blade outer air seal. 
     SUMMARY OF THE INVENTION 
     In the disclosed embodiment of this invention, a blade outer air seal is provided with a cooling channels that utilizes at least a plurality of distinct cooling schemes. In the disclosed embodiment, all of the cooling schemes utilized across the blade outer air seal are of the compact heat exchanger type. Of course, other type cooling schemes, such as the prior art ( FIG. 1 ) scheme formed by ceramic casting technology, can be utilized. In one embodiment, there are cooling schemes utilized adjacent the trailing edge of the blade outer air seal which will result in a relatively great pressure drop. The cooling schemes vary to decrease this pressure drop, moving in a direction towards the leading edge. As mentioned, the pressure ratio is greater at the trailing edge, and a higher pressure drop is acceptable. 
     As an example, one type of a cooling scheme which might be utilized adjacent the trailing edge includes a plurality of tortuous paths, and extends through a relatively long distance measured in a direction from the trailing edge to the leading edge. Air enters through passages at an outer peripheral surface of a body of the seal, passes through the tortuous path, and exits through exits at the inner periphery of the seal body. Similar “tortuous path” cooling schemes are utilized spaced from this first cooling scheme in a direction toward the leading edge, however, the spaced cooling schemes extend for a lesser distance such that the overall pressure drop decreases. 
     In the disclosed embodiment, and adjacent the leading edge, a distinct type cooling scheme is utilized wherein the tortuous paths are replaced by a plurality of pedestals within an open space. The pedestals increase the heat transfer surface area, but do not result in as much pressure drop as the tortuous path type cooling schemes mentioned above. 
     As known, typically, dozens of blade outer air seal sections are placed together circumferentially adjacent to other blade outer air seal sections. A cooling scheme is utilized adjacent one lateral edge of each section of blade outer air seal to provide cooling air at a relatively high pressure into a gap between adjacent sections. The cooling air supplied into the gap provides purge air to resist leakage of the products of combustion through this gap. 
     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  shows a portion of a prior art gas turbine engine. 
         FIG. 2  is a plan view of a number of cooling schemes within an example blade outer air seal. 
         FIG. 3  is a cross-sectional view along a portion of the  FIG. 2  scheme. 
         FIG. 4  is an enlarged portion of  FIG. 2 , along the circle  4 . 
         FIG. 5  shows a lost core for forming the various cooling schemes illustrated in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       FIG. 1  shows a portion of a gas turbine engine  20  having rotating turbine blades  22 , and a blade outer air seal  24  spaced slightly radially outwardly of the outermost portion of the turbine blade  22 . As shown, hooks  26  hold the blade outer air seal  24  into a housing  27 . As known, typically, dozens of sections of the blade outer air seal  24  are positioned circumferentially adjacent to each other to surround the turbine blades  22  and their rotor. 
     An air space  28  supplies air to a plurality of cooling channels  30  formed within a body of the blade outer air seal  24 . In general, these cooling channels  30  have been relatively thick in a radially outwardly extending dimension. Further, only one type of cooling scheme has been utilized throughout the blade outer air seal. As mentioned above, the cooling challenges and the fluid dynamics faced by the cooling air change as one moves from a leading edge of the blade outer air seal  24  toward a trailing edge (from left to right in  FIG. 1 ). 
       FIG. 2  is a cross-section through an inventive blade outer air seal section  50  having a leading edge  149  and a trailing edge  147 . Sides  145  and  143  sit adjacent to another section of blade outer air seal  50  when the blade outer air seal is assembled within a gas turbine engine. As shown in this figure, there are nine distinct internal cooling passages within the blade outer air seal  50 . 
     A first cooling scheme is provided by section  52 . Section  52  has inlet ports  54  that extend to a radially outer surface on the blade outer air seal body  50 . The cooling air passes into the inlets  54 , into an enlarged open space  55 , and over pedestals  58  before passing outwardly through outlets  56  in the side wall  143 . The pedestal type cooling schemes result in a relatively low pressure drop, and thus relatively high pressure air will be exiting the outlets  56  and into the gap between this blade outer air seal section  50  and an adjacent one. In this manner, the relatively high pressure air will purge leakage air away from the gap. The pedestals, as known, increase the heat transfer cross-sectional area and turbulence to provide more efficient and effective cooling. The section  52  is a compact heat exchanger section that is formed to be very thin in a radially outer dimension (into the plane of  FIG. 2 ). In this manner, relatively small cooling sections can be provided and can be tailored to the individual challenges of a particular area on the blade outer air seal  50 . 
     Another section  60  is spaced toward the leading edge  149  from the section  52 . Section  60  is configured to be much like section  52 , however, as can be appreciated, the gap between pedestals  58  is enlarged toward the leading edge, as such, the pressure drop is made to be less as one moves closer to the leading edge. 
     Another section  62  is formed adjacent the trailing edge. Section  62  is supplied with cooling air from inlets  64 , and that cooling air passes through a tortuous path around elongated strips  168 , and outwardly of outlets  66  in an inner peripheral surface of the blade outer air seal body  50 . This cooling air passes into the flow path of the products of combustion passing over the turbine. 
     As can be appreciated from  FIG. 3 , the inlet  64  extends to the outer periphery, the air passes over the strips  168 , and out of the outlet  66 . 
     Another cooling air section  68  receives air from an inlet  70 , passes air over elongated strips  74 , and outwardly through the outlet  75 . Another section  76  has inlet  78 , strips  82 , and outlet  80 . Yet another section  86  has inlet  88 , strips  190  and outlet  192 . 
     As can be appreciated from  FIG. 2 , the length of the sections  62 ,  68 ,  76  and  86  decreases as one moves from the trailing edge  147  towards the leading edge  149 . Again, this is because it would be desirable to reduce the overall pressure drop since the air must exit closer to the leading edge where the pressure ratio is lower. 
     As shown in  FIG. 4 , each of these cooling scheme sections provide a tortuous path with the air having to pass around the elongated strips. 
     Another cooling air section  90  is positioned adjacent the side  143 , and at the leading edge  149 . Section  90  has inlets  92 , and delivers through an open space over pedestals  98 , and outwardly through side outlets  96 , and forward outlets  94 . Side outlets  96  extend to the side  143 , whereas forward outlets  94  extend to the inner peripheral surface of the blade outer air seal body  50 . 
     Another section  100  has inlets  102 , outlets  104 , and pedestals  106 . Yet another section  108  has inlets  110 , side outlets  112 , forward outlets  114 , and pedestals  116 . Sections  90 ,  100  and  108  are all of the low pressure drop pedestal type, and thus do not reduce the pressure drop of the cooling air to a great extent such that it can exit into the working air, or the products of combustion. 
     A designer of a blade outer air seal can take advantage of the power provided by this invention to individually tailor cooling sections for the challenges faced by the particular area on a blade outer air seal. By utilizing this plurality of distinct type cooling schemes, the present invention provides more efficient and effective cooling. 
     The compact heat exchangers disclosed in this invention may be formed by a lost core mold technique. A core body is shown in  FIG. 5 .  FIG. 5  can also assist one in appreciating aspects of the shapes of the inlets and outlets, which may not be readily understandable from the plan view of  FIG. 2 . 
     It should be appreciated that  FIG. 5  actually shows a “mirror” of the cooling passages of  FIG. 2 . What  FIG. 5  shows is a core that will be put within a mold for forming the blade outer air seal. Once material has formed around this core, the core may be leached out of the material for forming the body, leaving cavities to provide the cooling air passages.  FIG. 5  includes reference numerals which are identical to those shown in  FIG. 2 , even though what is actually shown in  FIG. 5  is this core rather than the actual cooling passages. 
     Although a preferred 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.