Abstract:
Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, a representative blade outer air seal assembly for a gas turbine engine includes: an annular arrangement of outer air seal segments defining an inner diameter surface; intersegment gaps located between the outer air seal segments, each of the gaps being located between a corresponding adjacent pair of the segments; and recesses spaced about the inner diameter surface, each of the recesses communicating with a corresponding one of the gaps.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH AND DEVELOPMENT 
     The U.S. Government may have an interest in the subject matter of this disclosure as provided for by the terms of contract number N00019-02-C-3003, awarded by the United States Navy, and contract number F33615-03-D-2345 DO-0009, awarded by the United States Air Force. 
    
    
     BACKGROUND 
     1. Technical Field 
     The disclosure generally relates to gas turbine engines. 
     2. Description of the Related Art 
     A typical gas turbine engine incorporates a compressor section and a turbine section, each of which includes rotatable blades and stationary vanes. Within a surrounding engine casing, the radial outermost tips of the blades are positioned in close proximity to outer air seals. Outer air seals are parts of shroud assemblies mounted within the engine casing. Each outer air seal typically incorporates multiple segments that are annularly arranged within the engine casing, with the inner diameter surfaces of the segments being located closest to the blade tips. 
     SUMMARY 
     Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, an exemplary embodiment of a blade outer air seal assembly for a gas turbine engine comprises: an annular arrangement of outer air seal segments defining an inner diameter surface; intersegment gaps located between the outer air seal segments, each of the gaps being located between a corresponding adjacent pair of the segments; and recesses spaced about the inner diameter surface, each of the recesses communicating with a corresponding one of the gaps. 
     An exemplary embodiment of a gas turbine engine comprises: a compressor; a combustion section; a turbine operative to drive the compressor responsive to energy imparted thereto by the combustion section, the turbine having a rotatable set of blades; and a blade outer air seal assembly positioned radially outboard of the blades, the outer air seal assembly having an annular arrangement of outer air seal segments, intersegment gaps and recesses, the outer air seal segments defining an inner diameter surface, the intersegment gaps being located between the outer air seal segments, each of the gaps being located between a corresponding adjacent pair of the segments, the recesses being spaced about the inner diameter surface, and each of the recesses communicating with a corresponding one of the gaps. 
     An exemplary embodiment of a blade outer air seal segment comprises: a blade arrival end; a blade departure end; and an inner diameter surface extending at least partially between the blade arrival end and the blade departure end, at least a portion of the inner diameter surface being arcuately shaped as defined by a radius of curvature, a radially innermost portion of the inner diameter surface in a vicinity of the blade arrival end being located outboard of the radius of curvature. 
     Other systems, methods, features and/or advantages of this disclosure will be or may become apparent to one with skill in the art upon examination of the following drawings and detailed description. It is intended that all such additional systems, methods, features and/or advantages be included within this description and be within the scope of the present disclosure. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Many aspects of the disclosure can be better understood with reference to the following drawings. The components in the drawings are not necessarily to scale. Moreover, in the drawings, like reference numerals designate corresponding parts throughout the several views. 
         FIG. 1  is a schematic diagram depicting an exemplary embodiment of a gas turbine engine. 
         FIG. 2  is a partially cut-away, schematic diagram depicting a portion of the embodiment of  FIG. 1 . 
         FIG. 3  is a partially cut-away, schematic diagram depicting a portion of the shroud assembly of the embodiment of  FIGS. 1 and 2 . 
         FIG. 4  is a partially cut-away, schematic diagram depicting a portion of another embodiment of a blade outer air seal. 
         FIG. 5  is a partially cut-away, schematic diagram depicting a portion of another embodiment of a blade outer air seal. 
         FIG. 6  is a partially cut-away, schematic diagram depicting a portion of another embodiment of a blade outer air seal. 
     
    
    
     DETAILED DESCRIPTION 
     Gas turbine engines and related systems involving blade outer air seals are provided, several exemplary embodiments of which will be described in detail. In some embodiments, outer air seal segments incorporate blade arrival portions that include surfaces located radially outboard of corresponding surfaces of blade departure portions of adjacent segments. Thus, a spaced arrangement of recesses is provided about the inner diameter surface defined by the segments. Configuring the surfaces of the blade arrival portions in such a manner may tend to reduce wear of those surfaces. 
     Referring now in more detail to the drawings,  FIG. 1  is a schematic diagram depicting an exemplary embodiment of a gas turbine engine. As shown in  FIG. 1 , engine  100  incorporates a fan  102 , a compressor section  104 , a combustion section  106  and a turbine section  108 . Various components of the engine are housed within an engine casing  110 , such as a blade  112  of the low-pressure turbine, that extends along a longitudinal axis  114 . Although engine  100  is configured as a turbofan engine, there is no intention to limit the concepts described herein to use with turbofan engines as various other configurations of gas turbine engines can be used. 
     A portion of engine  100  is depicted in greater detail in the schematic diagram of  FIG. 2 . In particular,  FIG. 2  depicts a portion of blade  112  and a corresponding portion of a shroud assembly  120  that are located within engine casing  110 . Notably, blade  112  is positioned between vanes  122  and  124 , detail of which has been omitted from  FIG. 2  for ease of illustration and description. 
     As shown in  FIG. 2 , shroud assembly  120  is positioned between the rotating blades and the casing. The shroud assembly generally includes an annular mounting ring  123  and an annular outer air seal  125  attached to the mounting ring and positioned adjacent to the blades. Various other seals are provided both forward and aft of the shroud assembly. However, these various seals are not relevant to this discussion. 
     Attachment of the outer air seal to the mounting ring in the embodiment of  FIG. 2  is facilitated by interlocking flanges. Specifically, the mounting ring includes flanges (e.g., flange  126 ) that engage corresponding flanges (e.g., flange  128 ) of the outer air seal. Other attachment techniques may be used in other embodiments. 
     With respect to the annular configuration of the outer air seal, outer air seal  125  is formed of multiple arcuate segments, portions of two of which are depicted schematically in  FIG. 3 . As shown in  FIG. 3 , adjacent segments  140 ,  142  of the outer air seal are oriented in an end-to-end relationship, with an intersegment gap  150  located between the segments. 
     Portions defining the intersegment gap include a blade departure end  152  of segment  140  and a blade arrival end  154  of segment  142 . Generally, the ends interlock with each other with the intersegment gap varying in shape between embodiments. 
     A recess  160 , which communicates with the gap, also is defined by at least a portion of one of the ends. In the embodiment of  FIG. 3 , the recess is defined by a surface of segment  142 . Specifically, a portion  162  of an inner diameter surface of segment  142  is located at a distance R 1  from the longitudinal axis ( 114 ) of the engine and a portion  164  of the inner diameter surface is located at a greater distance from the longitudinal axis, i.e., located up to a distance R 2  from the longitudinal axis. Notably, portion  164  defines the recess, i.e., R 2  is longer than R 1 . It should also be noted that portion  162  of the embodiment of  FIG. 3  extends to the blade departure end of segment  142  (not shown) such that the inner diameter surface  166  of the blade departure end is located at distance R 1 . Since segment  140  and  142  are duplicate components in this embodiment, the inner diameter surface of blade departure end  152  of segment  140  is located at distance R 1 . Thus, the inner diameter surface in the vicinity of the blade arrival end is positioned radially outboard of the inner diameter surface in the vicinity of the blade departure end of the adjacent segment. Stated differently, a radially innermost portion of the blade arrival end is located radially outboard of a radially innermost portion of the blade departure end 
     The aforementioned configuration may tend to reduce stresses and corresponding wear exhibited by the blade arrival end over time. Notably, the advancing suction side of each rotating blade (e.g., side  170  of blade  112 ) tends to promote a radial inboard-directed ingestion flow of hot gas (depicted by the solid arrow) from the intersegment gap. In contrast, the retreating pressure side of each rotating blade (e.g., side  172  of blade  112 ) tends to promote a radial outboard-directed ingestion flow of hot gas (depicted by the dashed arrow) into the intersegment gap. By ensuring that a portion of the blade arrival end of a segment is located radially outboard of a corresponding portion of the blade departure end of an adjacent segment, a pressure dam condition can be avoided that can result in pressure augmentation experienced by the inner diameter surface at the blade arrival end. Such pressure augmentation can result in increased hot gas ingestion into the intersegment gap, which can lead to component deterioration. 
     Additionally or alternatively, ensuring that a portion of the blade arrival end of a segment is located radially outboard of a corresponding portion of the blade departure end of an adjacent segment may prevent an augmented heat transfer coefficient and heat load at the blade arrival end. Notably, avoiding such an augmented heat transfer coefficient and heat load could retard segment erosion at the blade arrival end. 
     Locating an inner diameter surface of a blade arrival end outboard of a corresponding surface of a blade departure end can be accomplished in a variety of manners. By way of example, the embodiment of  FIG. 3  uses a portion  164  of the inner diameter surface that is arcuately shaped. Specifically, portion  164  exhibits an outside radius curvature. In other embodiments, a different curvature (e.g., outside radius or compound curves) or no curvature (e.g., a planar surface) can be used. 
     In contrast, the embodiment of  FIG. 4  involves an inner diameter surface that exhibits an inside radius curvature. In particular, segments  180 ,  182  are oriented in an end-to-end relationship, with an intersegment gap  184  located between the segments. Portions defining the intersegment gap include a blade departure end  186  of segment  180  and a blade arrival end  188  of segment  182 . 
     A recess  190  communicates with gap  184  that is defined by portion  192  of the inner diameter surface of segment  182 . As shown in  FIG. 4 , portion  192  exhibits an inside radius curvature. Notably, the corresponding portion  194  of segment  180  does not exhibit a curvature. 
     Another embodiment is depicted schematically in  FIG. 5 . As shown in  FIG. 5 , adjacent segments  210 ,  212  are oriented in an end-to-end relationship, with an intersegment gap  214  located between the segments. Portions defining the intersegment gap include a blade departure end  216  of segment  210  and a blade arrival end  218  of segment  212 . 
     A recess  220  communicates with gap  214  that is defined by portion  222  of the inner diameter surface of segment  212 . As shown in  FIG. 5 , portion  222  exhibits an inside radius curvature, Notably, surface  226  of the blade departure end exhibits an outside radius curvature that complements the contour of portion  222  of segment  212 . 
     Another embodiment is depicted schematically in  FIG. 6 . As shown in  FIG. 6 , adjacent segments  230 ,  232  are oriented in an end-to-end relationship, with an intersegment gap  234  located between the segments. Portions defining the intersegment gap include a blade departure end  236  of segment  230  and a blade arrival end  238  of segment  232 . 
     A recess  240  communicates with gap  234  that is defined by portion  242  of the inner diameter surface of segment  232  and portion  244  of the inner diameter surface of segment  230 . As shown in  FIG. 6 , portion  242  exhibits an inside radius curvature, and portion  244  of the inner diameter surface of segment  230  exhibits an inside radius curvature. Notably, however, surface  246  of the blade departure end exhibits an outside radius curvature that complements the contour of portion  242  of segment  232 . 
     It should be emphasized that the above-described embodiments are merely possible examples of implementations set forth for a clear understanding of the principles of this disclosure. Many variations and modifications may be made to the above-described embodiments without departing substantially from the spirit and principles of the disclosure. All such modifications and variations are intended to be included herein within the scope of this disclosure and protected by the accompanying claims.