Patent Publication Number: US-11384654-B2

Title: Mateface for blade outer air seals in a gas turbine engine

Description:
TECHNICAL FIELD 
     The present disclosure relates generally to gas turbine engine flowpath construction, and more specifically to a contoured mateface configuration for utilization in a blade outer air seal. 
     BACKGROUND 
     Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some examples, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine as well. 
     Each of the compressor, combustor and turbine section are fluidly connected via a primary flowpath, with the outer diameter of the primary flowpath being defined at least partially by a set of circumferentially arranged blade outer air seals. 
     SUMMARY OF THE INVENTION 
     In one example, a gas turbine engine includes one of a turbine section and a compressor section comprised of a plurality of stages, at least one of the stages in the plurality of stages defining an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals, each blade outer air seal being spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap, wherein the mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals. 
     In another example of the above gas turbine engine, the mateface gap is defined by a first surface of a first blade outer air seal in the plurality of circumferentially arranged blade outer air seals and defined by a second surface of a second blade outer air seal in the plurality of circumferentially arranged blade outer air seals. 
     In another example of the any of the above gas turbine engines, the first surface and the second surface are complimentary angled planar surfaces relative to the gas turbine engine radius. 
     In another example of the any of the above gas turbine engines, the first surface and the second surface are planar surfaces and are oblique to each other. 
     In another example of the any of the above gas turbine engines, at least one of the first surface and the second surface includes a curvature. 
     In another example of the any of the above gas turbine engines, each of the first surface and the second surface includes a curvature. 
     In another example of the any of the above gas turbine engines, the curvature of the first surface and the curvature of the second surface is complimentary. 
     In another example of the any of the above gas turbine engines, at least one of the curvatures is a complex curvature. 
     In another example of the any of the above gas turbine engines, each of the curvatures is a complex curvature. 
     In another example of the any of the above gas turbine engines, at least one of the mateface gaps includes a nozzle fluid exit. 
     In another example of the any of the above gas turbine engines, each of the mateface gaps includes a nozzle fluid exit. 
     In another example of the any of the above gas turbine engines, an outer diameter portion of each mateface gap is sealed via an interstage seal. 
     In another example of the any of the above gas turbine engines, each interstage seal is a featherseal. 
     One exemplary method for film cooling an internal surface of a gas turbine engine component includes angling a mateface gap such that air exiting the mateface gap is directed along an inner diameter surface of one of an outer diameter component. 
     In another example of the above described method, the outer diameter component is a blade outer air seal. 
     Another example of the above described method includes diffusing air exiting the mateface gap using non-complimentary blade outer air seal surfaces. 
     Another example of the above described method includes accelerating air exiting the mateface gap. 
     An exemplary blade outer air seal for a gas turbine engine includes a platform defining an inner diameter surface, a first mateface surface extending from a first end of the inner diameter surface and a second mateface surface extending from the inner diameter surface, at least one of the first mateface surface and the second mateface surface being oblique to a radius of a gas turbine engine. 
     In another example of the above blade outer air seal, the first mateface surface and the second mateface surface are complimentary. 
     In another example of any of the above blade outer air seals, at least one of the first mateface surface and the second mateface surface is curved. 
     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  illustrates an exemplary gas turbine engine according to a first example. 
         FIG. 2  schematically illustrates a partial view of an exemplary turbine stage. 
         FIG. 3  schematically illustrates a first example interstage gap. 
         FIG. 4  schematically illustrates a second example interstage gap. 
         FIG. 5  schematically illustrates a third example interstage gap. 
         FIG. 6  schematically illustrates a fourth example interstage gap. 
         FIG. 7  schematically illustrates a fifth example interstage gap. 
         FIG. 8  schematically illustrates an interstage gap including a nozzle feature. 
     
    
    
     DETAILED DESCRIPTION 
       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 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a housing  15  such as a fan case or nacelle, and also 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 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 a 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  may be 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 the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  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 and less than about 5:1. 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 (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—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 (350.5 meters/second). 
     Within the compressor section  24  and the turbine section  28  are multiple stages with each stage including a set of circumferentially arranged rotors paired with a set of circumferentially arranged stators. Radially outward of the rotors in a given stage is an outer diameter defined by circumferentially arranged blade outer air seals, with the blade outer air seals being connected via intersegment seals. In order to cool the blade outer air seals, air is flowed circumferentially through the blade outer air seal and expelled from a circumferential edge of the blade outer air seal, into a gap between the segments. The facing surfaces of adjacent blade outer air seals are referred to as matefaces. The air is ejected from between the matefaces into the core gaspath flow, and expelled along the core gaspath. Any excess cooling capacity in the ejected air is wasted and represents a loss of efficiency. 
     With continued reference to  FIG. 1 ,  FIG. 2  schematically illustrates a partial stage  100  including a rotor  110  and multiple circumferentially arranged blade outer air seals  120  radially outward of the rotor  110 . While only a single rotor  110  is illustrated for explanatory effect, a practical implementation will include substantially more rotors. Each of the blade outer air seals  120  is connected to the circumferentially adjacent blade outer air seal  120  via an intersegment seal  122 . The intersegment seal  122  seals a radially outward end of an intersegment gap  124  between adjacent blade outer air seals  120 . In one example, the intersegment seal  122  is a feather seal. An airstream  130  is provided to each blade outer air seal  120 , and is flowed circumferentially through the blade outer air seal  120  to provide cooling to the blade outer air seal  120 . 
     Each of the blade outer air seals includes two circumferentially facing sides  126 ,  128  with at least one of the sides  126 ,  128  being at least partially skew relative to the radius of the engine  20 . By skewing at least one of the sides  126 ,  128  the airflow passed into the intersegment gap  124  is directed along a radially inward facing surface  121  of the blade outer air seal  120 , thereby enhancing the cooling effect provided to the blade outer air seal  120 . 
     The directing and contouring of the sides  126 ,  128  serves a dual purpose of directing leakage or post-part-cooling flow such that the flow cools the facing surface  126 ,  128  and provides a film coverage of an adjacent blade outer air seal  120  before being disposed of in the primary flowpath. In addition to this, the contours of the facing surfaces  126 ,  128  are in some examples configured to reduce ingestion of gaspath air into the intersegment cavity  124 , thereby reducing oxidization of the blade outer air seal  120 . 
     In order to further enhance the film cooling effect, and prevent flowpath air from being drawn into the interstage gap  124 , each interstage gap  124  is angled aligned with the rotation  102  direction of the rotors  110  such that the outlet of the interstage gap  124  is rotationally after the inlet (i.e. the outlet is counterclockwise of the inlet for a system with counterclockwise rotor rotation). In alternative examples, the rotation  102  can be reversed with the same alignment of the interstage gaps  124 . 
     With continued reference to  FIGS. 1 and 2 ,  FIGS. 3 through 7  illustrate potential facing surfaces pairs  302  (i.e. matefaces) defining an interstage gap  306  that could be used within a gas turbine engine. While illustrated as isolated surface pairs  302 , it is appreciated that exemplary gas turbine engines could be constructed using a single interstage gap configuration in between each blade outer air seal, or any combination of the illustrated interstage gap configurations within a given stage. In each example, the configuration assumes a clockwise rotating stage. In examples where the stage is rotating counterclockwise, the orientation of each of the exemplary interstage gaps  124  would be mirrored. 
     With regards to the first example facing surface pairs  302 , illustrated in  FIG. 3 , each of the facing surfaces is straight and angled complimentary to the paired surface. The straight surfaces reduce diffusion of the air exiting the interstage gap  306 , and the complimentary angles direct the air along the interior surface  304  of the counterclockwise blade outer air seal. 
     With regards to the second example facing surface pair  302 ,  FIG. 4  illustrates an example where both facing surfaces  301 ,  303  are angled, and where a counterclockwise surface  301  is positioned at a steeper angle than the clockwise surface  303 . The straight surfaces  301 ,  303  minimize diffusion, while the steeper counterclockwise surface  301  directs air exiting the interstage gap  306  along the counterclockwise surface  304  to enhance film cooling. 
     With regards to the third example facing surface pair  302 ,  FIG. 5  illustrates an example where the counterclockwise surface  301  is curved, and the clockwise surface  303  includes a complementary curve with a diffuser portion  305 . The example of  FIG. 5  utilizes the complimentary curvature to direct airflow exiting the interstage gap  306  along the inward surface  304  of the counterclockwise blade outer air seal. The diffuser portion  305  is a rounded end of the clockwise surface  303 . The diffuser portion  305  introduces diffusion into the air exiting the interstage gap  306 , while still maintaining the general direction of airflow toward the counterclockwise surface  304 . In addition, the diffuser portion  305  provides a sufficient cross section to reduce a cooling requirement of the diffuser portion  305 . 
     With regards to the fourth example facing surface pair  302 , illustrated in  FIG. 6 , each of the surfaces  301 ,  303  includes a complimentary complex curvature. As used herein, the complex curvature refers to the curvatures of the surfaces  301 ,  303  including multiple turns with arcs having an inconsistent radius of curvature. The utilization of complex curvatures can provide better or more efficient film coverture, outboard sealing, more consistent diffusion, better structural properties, better mateface cooling, as well as other benefits relative to other example architectures. 
     With regards to the fifth example facing surface pair  302 , the facing surface pair  302  of  FIG. 7  is a hybrid with the counterclockwise surface  301  being curved, and the clockwise surface  303  being straight. As illustrated herein, the clockwise surface  303  is aligned with the radius of the engine, however it is contemplated that alternative examples where the surface is angled relative to the radius (as in the examples of  FIGS. 3 and 4 ). 
     With continued reference to  FIGS. 1-7 ,  FIG. 8  schematically illustrates an interstage gap  402  including a cross sectional outlet area  410  that is smaller than a cross sectional area  412  of the gap  402  where the air enters the gap  402  from the blade outer air seals  404 ,  406 . The creation of the smaller cross section area  410  induces a nozzle effect accelerating the airflow as it is directed to the counterclockwise surface  408 . The acceleration allows for the air to be targeted and provide more efficient cooling in some examples. While illustrated with regards to paired straight surfaces it is appreciated that any of the configurations utilizing straight surfaces, contoured surfaces and/or complex curvatures could incorporate the nozzle feature in a similar manner. 
     While described above and illustrated in  FIGS. 2-8  within the context of turbine section blade outer air seals, it is appreciated that the contoured paired faces can be beneficially incorporated into any circumferentially arranged flowpath boundary component and are not limited in application to blade outer air seals or limited to turbine stages. By way of example, the contoured matefaces could be applied to static vanes, structural supports, and the like and/or may be disposed in the compressor section of the gas turbine 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.