Patent Publication Number: US-11661865-B2

Title: Gas turbine engine component

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This disclosure is a continuation of is U.S. patent application Ser. No. 17/329,510 filed May 25, 2021, which is a division of U.S. patent Ser. No. 16/019,972 filed Jun. 27, 2018, which is now U.S. Pat. No. 11,022,002 granted Jun. 1, 2021. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. 
     The efficiency of the engine is impacted by ensuring that the products of combustion pass in as high a percentage as possible across the turbine blades. Leakage around the blades reduces efficiency. 
     Thus, a blade outer air seal is provided radially outward of the blades to prevent leakage radially outwardly of the blades. The blade outer air seal may be held radially outboard from the rotating blade via connections on the case or a blade outer air seal support structure. The clearance between the blade outer air seal and a radially outer part of the blade is referred to as a tip clearance. Maintaining a proper tip clearance improves the efficiency of the gas turbine engine by reducing the amount of air leaking past the blade tips. 
     SUMMARY 
     In one exemplary embodiment, an attachment body for a blade outer air seal includes a leading edge connected to a trailing edge by a radially inner wall and a radially outer wall. At least one forward hook extends from the radially outer wall. At least one aft hook extends from the radially outer wall. At least one post extends from the radially outer surface and has a blade outer air seal (BOAS) guide surface. 
     In a further embodiment of any of the above, the radially outer surface includes at least one BOAS attachment surface. 
     In a further embodiment of any of the above, at least one BOAS attachment surface includes a pair BOAS attachment surfaces each located adjacent an opposing circumferential side of the attachment body. 
     In a further embodiment of any of the above, each of the pair of BOAS attachment surfaces define an arced surface. 
     In a further embodiment of any of the above, the arced surface includes a varying radius of curvature in an axial direction. 
     In a further embodiment of any of the above, the arced surface includes a constant radius of curvature in the axial direction. 
     In a further embodiment of any of the above, at least one post includes a pair of posts each having the BOAS guide surface facing a circumferential side of the attachment body. 
     In a further embodiment of any of the above, at least one aft hook includes a pair of aft hooks each including an anti-rotation tab. 
     In a further embodiment of any of the above, at least one post includes a pair of posts each having the BOAS guide surface facing a circumferential side of the attachment body. 
     In another exemplary embodiment, a seal assembly includes at least one blade outer air seal (BOAS) which includes a base portion that extends between a leading edge and a trailing edge. A forward wall and an aft wall extend radially outward from the base portion to a radially outer portion. The radially outer portion is spaced from the base portion and at least partially defines a passage with the forward wall, aft wall, and base portion. At least one attachment body is located at least partially within the passage. 
     In a further embodiment of any of the above, the attachment body includes a radially outer surface that has at least one post with a BOAS guide surface. 
     In a further embodiment of any of the above, the radially outer surface includes a BOAS attachment surface in contact with at least one of the blade outer air seals. 
     In a further embodiment of any of the above, the radially outer surface includes a pair of BOAS attachment surfaces each in contact with a corresponding one of a first BOAS and a second BOAS. 
     In a further embodiment of any of the above, each of the pair of BOAS attachment surfaces define an arced surface. 
     In a further embodiment of any of the above, at least one post includes a pair of posts each having the BOAS guide surface facing a circumferential side of the attachment body. 
     In a further embodiment of any of the above, the attachment body includes a pair of aft hooks each including an anti-rotation tab. 
     In another exemplary embodiment, a method of assembling a blade outer air seal assembly comprising the steps of engaging a first blade outer air seal (BOAS) with a first attachment surface on a first attachment body. A second BOAS is engaged with a second attachment surface on the first attachment body. The attachment body prevents rotation relative to the first BOAS with a first post and the second BOAS with a second post. 
     In a further embodiment of any of the above, the attachment body includes a radially outer surface and the first post and the second post are located on the radially outer surface, the first post includes a first BOAS guide surface and the second post includes a second BOAS guide surface. 
     In a further embodiment of any of the above, the first attachment surface and the second attachment surface each define an arced surface. 
     In a further embodiment of any of the above, anti-rotating the attachment body relative to an engine static structure with at least one tab that extends from an aft hook on the attachment body. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a schematic view of an example gas turbine engine according to a non-limiting example. 
         FIG.  2    is an enlarged schematic view of a portion of a turbine section. 
         FIG.  3    is perspective view of a blade outer air seal. 
         FIG.  4    is a side view of the blade outer air seal. 
         FIG.  5    is a perspective view of an attachment body. 
         FIG.  6    is a partially assembled view of the blade outer air seal and attachment body of  FIGS.  3  and  5   . 
         FIG.  7    is a perspective view of the pair of blade outer air seals of  FIG.  6    assembled with the attachment body of  FIG.  5   . 
         FIG.  8    is a cross-sectional view along line  8 - 8  of  FIG.  7   . 
         FIG.  9    schematically illustrates multiple blade outer air seals from  FIG.  3    arranged into a segmented ring. 
     
    
    
     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 nacelle  15 , 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 1 bm of fuel being burned divided by 1 bf 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). 
       FIG.  2    illustrates an enlarged schematic view of the high pressure turbine  54 , however, other sections of the gas turbine engine  20  could benefit from this disclosure, such as the compressor section  24  or low pressure turbine  46 . In the illustrated example, the high pressure turbine  54  includes a one-stage turbine section including a first rotor assembly  60 . In another example, the high pressure turbine  54  could include a two-stage high pressure turbine section with multiple rotor assemblies separated by stators. 
     The first rotor assembly  60  includes a plurality of first rotor blades  62  circumferentially spaced around a first disk  64  to form an array. Each of the plurality of first rotor blades  62  include a first root portion  72 , a first platform  76 , and a first airfoil  80 . Each of the first root portions  72  is received within a respective first rim  66  of the first disk  64 . The first airfoil  80  extends radially outward toward a blade outer air seal (BOAS)  82 . The BOAS  82  is attached to the engine static structure  36  by an attachment body  84  engaging retention hooks  86  on the engine static structure  36 . In the illustrated example, the attachment body  84  is a separate structure from the BOAS  82  and the engine static structure  36  shown in  FIG.  2    could be a portion of an engine case or a support structure. 
     The plurality of first rotor blades  62  are disposed in the core flow path C that is pressurized in the compressor section  24  then heated to a working temperature in the combustor section  26 . The first platform  76  separates a gas path side inclusive of the first airfoils  80  and a non-gas path side inclusive of the first root portion  72 . 
     A plurality of vanes  90  are located axially upstream of the plurality of first rotor blades  62 . Each of the plurality of vanes  90  includes at least one airfoil  92  that extends between a respective vane inner platform  94  and a vane outer platform  96 . In another example, each of the array of vanes  90  include at least two airfoils  92  forming a vane double. 
     As shown in  FIGS.  3  and  4   , the blade outer air seal  82  includes a leading edge  98  and a trailing edge  100 . In the illustrated example, the BOAS  82  is made of a ceramic matrix composite (CMC) and includes a forward wall  102  and an aft wall  104  that extend radially outward from a base portion  108  to an outer wall  106 . The BOAS  82  may also be made of a monolithic ceramic. The base portion  108  extends between the leading edge  98  and the trailing edge  100  and defines a gas path on a radially inner side and a non-gas path on a radially outer side. The outer wall  106  includes a generally constant thickness and constant position in the radial direction such that an outer surface of the outer wall  106  is planer. In this disclosure, forward, aft, upstream, downstream, axial, radial, or circumferential is in relation to the engine axis A unless stated otherwise. 
     In the illustrated example, circumferentially outward of the outer wall  106 , the forward wall  102  extends a distance D 1  from a radially inner edge of the BOAS  82  and the aft wall  104  extends a distance D 2  from the radially inner edge of the BOAS  82  with the distance D 2  being greater than the distance D 1 . By having the distance D 1  being less than the distance D 2 , the BOAS  82  can be assembled into a ring (see  FIG.  9   ) with multiple blade outer air seals  82  and have a greater amount of clearance along a leading region for assembly into the gas turbine engine  20 . Assembly time of the gas turbine engine can be reduced when the ring of blade outer air seals  82  does not need to be installed individually but as a continuous ring with multiple segments (See  FIG.  9   ). 
     The forward wall  102 , the aft wall  104 , the outer wall  106 , and the base portion  108  of the BOAS  82  define a passage  110  for accepting the attachment body  84 . A radially inner side of the base portion  108  at least partially defines the core flow path C and is located adjacent a tip of the first airfoil  80  (See  FIG.  2   ). 
       FIG.  5    illustrates the attachment body  84 . The attachment body  84  includes the base portion  108  extending between a leading edge  112  and a trailing edge  114 . The leading edge  112  and the trailing edge  114  are connected by a radially inner surface  116  and a radially outer surface  118 . Corners of the attachment body  84  includes notches to facilitate ease of installation into a corresponding one of the BOAS  82 . The radially inner surface  116  and the radially outer surface  118  also extend between opposing circumferential sides  120  on circumferential end portions of the attachment body  84 . The radially inner surface  116  can be a planer surface or an arced surface such that the radially inner surface is conical or includes a radius of curvature. 
     The radially outer surface  118  includes a perimeter portion  118 A that surrounds a recessed portion  118 B. The recessed portion  118 B includes a wall  119  that surrounds the recessed portion  118 B and connects the recessed portion  118 B to the perimeter portion  118 A. The perimeter portion  118 A includes a BOAS attachment surface  121  adjacent each of the circumferential sides  120  on circumferential end portions of the attachment body  84 . Each of the BOAS attachment surfaces  121  are located adjacent or in contact with one of the BOAS  82  as shown in  FIGS.  7  and  8   . At least one of the BOAS attachment surfaces  121  define an arced surface such that the BOAS attachment surface  121  includes a constant radius of curvature, such as with a cylinder, or a radius of curvature that varies in the axial direction defining a conical shape. 
     A forward hook  122  extends from the perimeter portion  118 A of the radially outer surface  118  of the attachment body  84  adjacent the leading edge  112 . The forward hook  122  includes a radially outward extending portion  122 A and an axially forward extending portion  122 B. Although only a single forward hook  122  is shown in the illustrated example of  FIG.  5   , more than one forward hook  122  could be incorporated into the attachment body  84 . In the illustrated example, the axially forward extending portion  122 B on the forward hook  122  engages at least one of the retention hooks  86  on the engine static structure  36  (See  FIG.  2   ). 
     At least one aft hook  124  also extends from the perimeter portion  118 A of the radially outer surface  118  and includes a portion extending radially outward  124 A and a portion  124 B extending axially forward and aft of the portion extending radially outward. The portion  124 B on each of the aft hooks  124  includes a tab  125  that extends axially forward. The tabs  125  engage the retention hooks  86  or the engine static structure  36  to provide an anti-rotation function to prevent or reduce the attachment body  84  from rotating relative to the retention hooks  86 /engine static structure  36  (See  FIG.  2   ). 
     A pair of posts  126  also extend from the radially outer surface  118 . The pair of posts  126  engage the BOAS  82  to prevent the BOAS  82  from rotating relative to the attachment body  84 . The pair of posts each include a BOAS guide surface  126 A. In the illustrated example, the BOAS guide surface  126 A contacts the BOAS  81  as shown in  FIGS.  7  and  8   . However, the BOAS guide surface  126 A could also be located in close proximity to the BOAS  81  or be spaced from the BOAS  81  by a wear liner. The pair of posts  126  includes an axial dimension that is greater than a circumferential dimension. In the illustrated example, the pair of posts  126  extend from the recessed portion  118 B of the radially outer surface  118  and the guide surface  126 A intersects the perimeter surface  118 A with a transition surface  126 B, such as a fillet or curved surface. 
       FIGS.  6 - 8    illustrate an assembly procedure for the BOAS  82  and attachment body  84 . As shown in  FIG.  6   , one of the attachment bodies  84  is radially and axially aligned with corresponding passages  110  in each of a pair of the BOAS  82 . The attachment body  84  can be moved circumferentially such that one of the circumferential sides  120  is accepting within the passage  110  in one of the BOAS  82 . Then the other BOAS  82  can be moved circumferentially until the other circumferential side  120  on the attachment body  84  is accepted within the passage  110  in the other BOAS  82 . Alternatively, the attachment body  84  can remain fixed while moving each of the pair of BOAS  82  circumferentially toward attachment body  84  until corresponding circumferential sides  120  are accepted within corresponding passages  110  in each of the BOAS  82 . The above procedures are continued until a plurality of BOAS  82  and attachment bodies  84  form a complete ring as shown in  FIG.  9   . 
     As shown in the cross-sectional view in  FIG.  8   , the guide surface  126 A of the posts  126  are located adjacent to or in direct contact with the outer wall  106  on the BOAS  82 . The posts  126  prevent the attachment body  84  from rotating relative to the BOAS  82 . The notches in the corners of the attachment body  84  as shown in  FIG.  5    also facilitate ease of insertion into the passages  110  by guiding the attachment body  84  into the passage  110  due to the reduce axial dimension of the attachment bodies  84  that result from the notches. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.