Patent Publication Number: US-2020300107-A1

Title: Cmc blade outer air seal

Description:
BACKGROUND 
     This application relates to a ceramic matrix composite blade outer air seal. 
     Gas turbine engines are known and typically include a compressor compressing air and delivering it into a combustor. The air is mixed with fuel in the combustor and ignited. Products of the combustion pass downstream over turbine rotors, driving them to rotate. 
     It is desirable to ensure that the bulk of the products of combustion pass over turbine blades on the turbine rotor. As such, it is known to provide blade outer air seals radially outwardly of the blades. Blade outer air seals have been proposed made of ceramic matrix composite fiber layers. 
     SUMMARY 
     In one exemplary embodiment, a blade outer air seal includes a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first wall is axially spaced from a second wall. The first and second walls extend from the base portion. An outer wall joins the first and second walls. The outer wall has a first edge and a second edge. Each of the edges have a first portion and a second portion arranged at a first angle relative to the first portion. 
     In a further embodiment of the above, the first and second circumferential edges are circumferentially inward of the first and second circumferential sides of the base portion. 
     In a further embodiment of any of the above, the first portion of the circumferential edges extends in a generally axial direction. 
     In a further embodiment of any of the above, the first angle is less than about 45°. 
     In a further embodiment of any of the above, the first angle is less than about 20°. 
     In a further embodiment of any of the above, the second portion is arranged axially forward of the first portion. A third portion is arranged axially aft of the first portion. The third portion is arranged at a second angle relative to the first portion. 
     In a further embodiment of any of the above, the second angle is smaller than the first angle. 
     In a further embodiment of any of the above, the base portion extends axially beyond the first wall. 
     In a further embodiment of any of the above, a slot extends through the outer wall. 
     In a further embodiment of any of the above, the first and second edges form mating surfaces configured to engage a support structure or carrier. 
     In a further embodiment of any of the above, the first wall, the second wall, and the outer wall have a same thickness. 
     In a further embodiment of any of the above, a film cooling hole extends through the base portion. 
     In a further embodiment of any of the above, the film cooling hole is between the first and second walls. 
     In a further embodiment of any of the above, the blade outer air seal is a ceramic matrix composite material. 
     In another exemplary embodiment, a turbine section for a gas turbine engine includes a turbine blade that extends radially outwardly to a radially outer tip and for rotation about an axis of rotation. A blade outer air seal has a plurality of segments mounted in a support structure via a carrier. The plurality of segments are arranged circumferentially about the axis of rotation and radially outward of the outer tip. Each segment has a first wall axially spaced from a second wall. The first and second walls are joined to a base portion and an outer wall. The outer wall has a first edge and a second edge. Each of the edges have a first portion and a second portion arranged at a first angle relative to the first portion. 
     In a further embodiment of any of the above, the first and second edges are engaged with the carrier. 
     In a further embodiment of any of the above, a wear liner is arranged within each segment. The wear liner has a radially extending tab engaged with the first portion. 
     In a further embodiment of any of the above, the base portion extends between first and second circumferential sides. The first and second circumferential edges are inward of first and second circumferential sides. 
     In a further embodiment of any of the above, the second portion is arranged axially forward of the first portion. A third portion is arranged axially aft of the first portion. The third portion is arranged at a second angle relative to the first portion. The second angle is smaller than the first angle. 
     In a further embodiment of any of the above, the blade outer air seal is a ceramic matrix composite material. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  shows an example turbine section. 
         FIG. 3  shows a portion of an exemplary blade outer air seal assembly. 
         FIG. 4  shows a cross section of the exemplary blade outer air seal assembly. 
         FIG. 5  shows an exemplary blade outer air seal. 
         FIG. 6  shows a side view of the exemplary blade outer air seal of  FIG. 5 . 
         FIG. 7  shows another embodiment of a blade outer air seal. 
         FIG. 8  shows a blade outer air seal assembly. 
     
    
    
     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 the 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 the exemplary gas turbine engine  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). 
       FIG. 2  shows a portion of an example turbine section  28 , which may be incorporated into a gas turbine engine such as the one shown in  FIG. 1 . However, it should be understood that other sections of the gas turbine engine  20  or other gas turbine engines, and even gas turbine engines not having a fan section at all, could benefit from this disclosure. 
     A turbine blade  102  has a radially outer tip  103  that is spaced from a blade outer air seal assembly  104  with a blade outer air seal (“BOAS”)  106 . The BOAS  106  may be made up of a plurality of seal segments  105  that are circumferentially arranged in an annulus about the central axis A of the engine  20 . The BOAS segments  105  may be monolithic bodies that are formed of a high thermal-resistance, low-toughness material, such as a ceramic matrix composite (“CMC”). The BOAS segments  105  are mounted to a BOAS support structure  110  via an intermediate carrier  112 . The support structure  110  may be mounted to an engine structure, such as engine static structure  36 . In some examples, the support structure  110  is integrated with engine static structure  36 . 
       FIG. 3  shows a portion of an example BOAS assembly  104 . The assembly  104  has a seal segment  105  with a carrier  112 . The carrier  112  may be segmented, with each segment arranged between adjacent seal segments  105 . The carrier  112  has a base portion  118  that is configured to engage with the seal segment  105 . In this example, an end of the base portion  118  fits within a passage  138  (shown in  FIG. 5 ) of the seal segment  105 . The carrier  112  has first and second hooks  114 ,  116  that extend radially outward from the base portion  118  for attaching the carrier  112  and seal segment  105  to the support structure  110 . The carrier  112  may have posts  119  that engage with an edge of the seal segment  105 , and help prevent rotation of the seal segment  105  relative to the carrier  112 . 
     A wear liner  162  may be arranged between the seal segment  105  and the carrier  112  in some examples. A feather seal  160  may be used for sealing between circumferential ends C 1 , C 2  of adjacent seal segments  105 . The feather seal  160  may extend along the axial length of the BOAS segment  105 . 
       FIG. 4  shows the BOAS assembly  104  with the support structure  110 . The support structure  110  has first and second hooks  115 ,  117  that extend radially inward and are configured to engage with the first and second hooks  114 ,  116  of the carrier  112 . In the illustrated embodiment, the hooks  114 ,  116  of the carrier  112  extend generally axially forward towards the leading edge  99 , while the hooks  115 ,  117  extend generally axially backwards towards the trailing edge  101 . However, the hooks  114 ,  116 ,  115 ,  117  may have different orientations, such as extending in the opposite direction, so long as the hooks  114 ,  116  of the carrier engage with the hooks  115 ,  117  of the support structure  110 . 
     The assembly  104  may include a front brush seal  164  and a dogbone or diamond seal  166  in some examples. These seals  164 ,  166  are engaged with the leading edge  99  of the BOAS segments  105 , and help maintain the axial position of the BOAS  106 . The seal  166  pushes the brush seal  164  axially forward and the BOAS segments  105  axially aft. 
       FIG. 5  illustrates an exemplary BOAS segment  105 . The seal segment  105  is a body that defines radially inner and outer sides R 1 , R 2 , respectively, first and second axial sides A 1 , A 2 , respectively, and first and second circumferential sides C 1 , C 2 , respectively. The radially inner side R 1  faces in a direction toward the engine central axis A. The radially inner side R 1  is thus the gas path side of the seal segment  105  that bounds a portion of the core flow path C. The first axial side A 1  faces in a forward direction toward the front of the engine  20  (i.e., toward the fan  42 ), and the second axial side A 2  faces in an aft direction toward the rear of the engine  20  (i.e., toward the exhaust end). That is, the first axial side A 1  corresponds to a leading edge  99 , and the second axial side A 2  corresponds to a trailing edge  101 . 
     In the illustrated example, the BOAS segment  105  includes a first axial wall  120  and a second axial wall  122  that extend radially outward from a base portion  124 . The first and second axial walls  120 ,  122  are axially spaced from one another. Each of the first and second axial walls  120 ,  122  extends along the base portion  124  in a generally circumferential direction along at least a portion of the seal segment  105 . The base portion  124  extends between the leading edge  99  and the trailing edge  101  and defines a gas path on a radially inner side and a non-gas path on a radially outer side. An outer wall  126  extends between the first and second axial walls  120 ,  122 . The outer wall  126  includes a generally constant thickness and constant position in the radial direction. The base portion  124 , first and second axial walls  120 ,  122 , and the outer wall  126  form a passage  138  that extends in a generally circumferential direction. In this disclosure, forward, aft, upstream, downstream, axial, radial, or circumferential is in relation to the engine axis A unless stated otherwise. The base portion  124  may extend axially forward and aft of the first and second walls  120 ,  122 , and provides a flat surface for sealing of the BOAS leading and trailing edges  99 ,  101 . For example, the base portion  124  includes a portion axially forward of the first axial wall  120  for engagement with seals  164 ,  166  (shown in  FIG. 4 ). 
     The outer wall  126  has first and second edges  130 ,  132 . The edges  130 ,  132  have tapered portions. A first portion  131 ,  133  of the edges  130 ,  132 , respectively, extends generally in the axial direction X. The first portions  131 ,  133  provide a flat face for engagement with the carrier  112 , and help prevent rotation of the seal segment  105  relative to the carrier  112 . Tapered portions upstream and downstream of the first portion  131 ,  133  are angled relative to the axial direction X. A second portion  134 ,  136  of the edges  130 ,  132 , respectively, is upstream of the first portions  131 ,  133 . The second portions  134 ,  136  are arranged at a first angle Θ 1  with respect to the first portions  131 ,  133 . A third portion  135 ,  137  of the edges  130 ,  132 , respectively, is downstream of the first portions  131 ,  133 . The third portions  135 ,  137  are arranged at a second angle Θ 2  with respect to the first portions  131 ,  133 . The second and third portions  134 ,  136 ,  135 ,  137  provide tapered faces, which may reduce stresses on the seal segment  105 . In one example embodiment, the first and second angles Θ 1 , Θ 2  are less than about 45° with respect to the axial direction X. In another embodiment, the first and second angles Θ 1 , Θ 2  are less than about 20° with respect to the axial direction X. The first angle Θ 1  may be greater than the second angle Θ 2 . In one example, the first angle Θ 1  is about 20° and the second angle Θ 2  is about 10°. 
     In the illustrated embodiment, the first portion  131 ,  133  is generally centered on the outer wall  126 . However, in other embodiments, the first portion  131 ,  133  may be moved axially forward or aft, depending on the carrier  112  and wear liner  162  to address varying torque loads. In one example embodiment, the first portion  131 ,  133  has a length in the axial direction of about 0.30 inches (7.62 mm). The axial length of the first portion  131 ,  133  provides a surface for mating with the carrier  112 . 
       FIG. 6  shows a side view of the BOAS segment  105 . The seal segment  105  may be formed of a ceramic matrix composite (“CMC”) material. Each seal segment  105  is formed of a plurality of CMC laminate plies  142 . The laminate plies  142  may be silicon carbide fibers, formed into a braided or woven fabric in each layer. The fibers may be coated by a boron nitride. In other examples, the BOAS segment  105  may be made of a monolithic ceramic. CMC components such as BOAS segments  105  are formed by laying fiber material, such as laminate sheets, in tooling, injecting a liquid resin into the tooling, and curing to form a solid composite component. The component may be densified by adding additional material to further stiffen the laminates. 
     Densification includes injecting material, such as a silicon carbide matrix material, into spaces between the fibers in the laminate plies. This may be utilized to provide 100% of the desired densification, or only some percentage. One hundred percent densification may be defined as the layers being completely saturated with the matrix and about the fibers. One hundred percent densification may be defined as the theoretical upper limit of layers being completely saturated with the matrix and about the fibers, such that no additional material may be deposited. In practice, 100% densification may be difficult to achieve. Although a CMC loop BOAS segment  105  is shown, other BOAS arrangements may be utilized within the scope of this disclosure 
     In an embodiment, the BOAS segment  105  is formed from two loops of CMC laminated plies. A first loop  144  comprises the inner-most layers relative to the respective passage  138 . A second loop  146  is formed about the first loop  144  to form the outermost layers relative to the passage  138 . In one example embodiment, the first and second loops  144 ,  146  are each formed from four laminated plies  142 . A noodle region  145  may be formed between the first and second loops  144 ,  146 . The noodle region  145  may be filled with a matrix material during densification, in some examples. In some examples, the base portion  124  may have additional reinforcement plies  143 . The reinforcement plies  143  may reduce the size of the noodle regions  145 , which strengthens the overall structure. 
     The transverse direction of the plies  142  helps evenly distribute stresses on the component. The shape of the seal segment  105  and the passage  138  allows for complex cooling arrangement and relatively low thermal stresses. The seal segment  105  also allows for multiple sealing surfaces and may accommodate different designs for the intermediate carrier  112 . The loop construction of the seal segment  105  also minimizes delamination when the seal segment  105  is secured to the support structure  110  via the carrier  112 . 
     In an example embodiment, the first wall  120 , second wall  122 , and outer wall  126  have a constant wall thickness of about 8 laminated plies  142 , with each plie  142  having a thickness of about 0.011 inches (0.279 mm). This structure may reduce thermal gradient stress. Although 8 laminated plies are described, BOAS constructed of more or fewer plies may fall within the scope of this disclosure. In one example, the first and second loops  144 ,  146  are formed from laminates wrapped around a core mandrel. In some embodiments, after the laminate plies  142  are formed into a seal segment  105 , additional features, such as edges  130 ,  132  are machined in to form mating surfaces and/or cooling holes. The seal segment  105  may be ultrasonically machined, for example. 
       FIG. 7  shows another example seal segment  205 . In this example, a hole  270  extends through the outer wall  226 . The hole  270  may be centered circumferentially on the seal segment  205 . In the illustrated example, the hole  270  is an elongated slot shape, however, other hole shapes may be used. The hole  270  allows cooling air to pass through the outer wall  226  into the passage  238  and through cooling holes  241 . The hole  270  may also provide weight reduction and may reduce stresses on the seal segment  205 . 
       FIG. 8  shows a BOAS assembly  104 . The assembly  104  includes a plurality of seal segments  105  and a plurality of carrier segments  112  arranged in an annulus about the engine axis A. There are an equal number of carrier segments  112  as seal segments  105 . In the illustrated example, there are  40  seal segments and  40  carrier segments. However, more or fewer segments may be used. 
     The disclosed BOAS arrangement reduces stress on the seal segment  105  by providing edges  130 ,  132  to engage with the carrier  112 . The edges  130 ,  132  have a flat portion  131 ,  133  to prevent rotation. The tapered portions  134 ,  135 ,  136 ,  137  of the edges  130 ,  132  reduce stresses on the seal segment  105 . The edges  130 ,  132  also permit tooling access for machining cooling holes  141  into the base portion  124 . The disclosed seal segment  105  permits cost effective manufacturing and assembly and allows the use of a ceramic BOAS. The ability to use a ceramic BOAS promotes a more stable assembly because ceramic materials are not as ductile as metallic materials. The disclosed CMC BOAS has simple features that are easily manufactured using CMC laminates. 
     In this disclosure, “generally axially” means a direction having a vector component in the axial direction that is greater than a vector component in the circumferential direction, “generally radially” means a direction having a vector component in the radial direction that is greater than a vector component in the axial direction and “generally circumferentially” means a direction having a vector component in the circumferential direction that is greater than a vector component in the axial direction. 
     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 disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.