Patent Publication Number: US-10323534-B2

Title: Blade outer air seal with cooling features

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This is a continuation of U.S. patent application Ser. No. 13/549,874, which was filed on Jul. 16, 2012. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine, and more particularly to a blade outer air seal (BOAS) that may be incorporated into a gas turbine engine. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     A casing of an engine static structure may include one or more blade outer air seals (BOAS) that provide an outer radial flow path boundary for the hot combustion gases. The BOAS surrounds rotor assemblies that carry one or more blades that rotate and extract energy from the hot combustion gases communicated through the gas turbine engine. The BOAS may be subjected to relatively extreme temperatures during gas turbine engine operation. 
     SUMMARY 
     A blade outer air seal (BOAS) for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. At least one cooling fin is disposed on the radially outer face between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of the foregoing BOAS, a plurality of cooling fins axially extend between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of either of the foregoing BOAS, at least one cooling fin extends across an entire length between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of any of the foregoing BOAS, at least one cooling fin axially extends between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of any of the foregoing BOAS, a plurality of cooling fins are circumferentially disposed about the radially outer surface of the seal body. 
     In a further non-limiting embodiment of any of the foregoing BOAS, the leading edge portion includes an engagement feature that receives a portion of a support structure of the gas turbine engine. 
     In a further non-limiting embodiment of any of the foregoing BOAS, a seal is attached to the radially inner face of the seal body. 
     In a further non-limiting embodiment of any of the foregoing BOAS, the seal is a honeycomb seal. 
     In a further non-limiting embodiment of any of the foregoing BOAS, a thermal barrier coating is applied to the radially inner face of the seal body between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of any of the foregoing BOAS, at least one cooling fin extends at a non-perpendicular angle relative to the radially outer face. 
     A gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor section. A blade outer air seal (BOAS) is associated with at least one of the compressor section and the turbine section. The BOAS includes a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion and at least one cooling fin disposed on the radially outer face between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of the foregoing gas turbine engine, the BOAS is positioned radially outward from a blade tip of a blade of at least one of the compressor section and the turbine section. 
     In a further non-limiting embodiment of either of the foregoing gas turbine engines, a plurality of cooling fins axially extend across the radially outer face between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, at least one cooling fin axially extends between the leading edge portion and the trailing edge portion. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, a plurality of cooling fins are disposed on the radially outer surface. A first portion of the plurality of cooling fins include a first length and a second portion of the plurality of cooling fins include a second length that is different from the first length. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, at least one cooling fin includes a first height adjacent to the leading edge portion and a second height that is different from the first height adjacent to the trailing edge portion. 
     A method of providing a blade outer air seal (BOAS) for a gas turbine engine, according to another exemplary aspect of the present disclosure includes, among other things, providing the BOAS with at least one cooling fin on a radially outer face of the BOAS. 
     In a further non-limiting embodiment of the foregoing method of providing a blade outer air seal (BOAS) for a gas turbine engine, the method may include a plurality of cooling fins circumferentially disposed about the radially outer face. 
     In a further non-limiting embodiment of either of the foregoing methods of providing a blade outer air seal (BOAS) for a gas turbine engine, the method communicates an airflow across the at least one cooling fin to cool the BOAS. 
     In a further non-limiting embodiment of any of the foregoing methods of providing a blade outer air seal (BOAS) for a gas turbine engine, the method may include providing at least one cooling fin extending axially between a leading edge portion and a trailing edge portion of the BOAS. 
     The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
         FIG. 2  illustrates a cross-section of a portion of a gas turbine engine. 
         FIG. 3  illustrates a perspective view of a blade outer air seal (BOAS). 
         FIG. 4  illustrates a portion of the BOAS of  FIG. 3 . 
         FIG. 5  illustrates another exemplary BOAS. 
         FIG. 6  illustrates exemplary cooling fins that can be incorporated into a BOAS. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The exemplary gas turbine engine  20  is a two-spool turbofan engine that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section  22  drives air along a bypass flow path B, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 . The hot combustion gases generated in the combustor section  26  are expanded through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, turboshaft engines. 
     The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine centerline longitudinal axis A. The low speed spool  30  and the high speed spool  32  may be mounted relative to an engine static structure  33  via several bearing systems  31 . It should be understood that additional bearing systems  31  may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  34  that interconnects a fan  36 , a low pressure compressor  38  and a low pressure turbine  39 . The high speed spool  32  includes an outer shaft  35  that interconnects a high pressure compressor  37  and a high pressure turbine  40 . In this embodiment, the inner shaft  34  and the outer shaft  35  are supported at various axial locations by bearing systems  31  positioned within the engine static structure  33 . 
     A combustor  42  is arranged between the high pressure compressor  37  and the high pressure turbine  40 . A mid-turbine frame  44  may be arranged generally between the high pressure turbine  40  and the low pressure turbine  39 . The mid-turbine frame  44  supports one or more bearing systems  31  of the turbine section  28 . The mid-turbine frame  44  may include one or more airfoils  46  that may be positioned within the core flow path C. 
     The inner shaft  34  and the outer shaft  35  are concentric and rotate via the bearing systems  31  about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor  38  and the high pressure compressor  37 , is mixed with fuel and burned in the combustor  42 , and is then expanded over the high pressure turbine  40  and the low pressure turbine  39 . The high pressure turbine  40  and the low pressure turbine  39  rotationally drive the respective high speed spool  32  and the low speed spool  30  in response to the expansion. 
       FIG. 2  illustrates a portion  100  of a gas turbine engine, such as the gas turbine engine  20  of  FIG. 1 . In this exemplary embodiment, the portion  100  represents part of the turbine section  28 . However, it should be understood that other portions of the gas turbine engine  20  could benefit from the teachings of this disclosure, including but not limited to, the compressor section  24 . 
     In this exemplary embodiment, a blade  50  (only one shown, although multiple blades could be circumferentially disposed about a rotor disk (not shown) within the portion  100 ) is mounted for rotation relative to a casing  52  of the engine static structure  33 . In the turbine section  28 , the blade  50  rotates to extract energy from the hot combustion gases that are communicated through the gas turbine engine  20 . The portion  100  can also include a vane assembly  54  supported within the casing  52  at a downstream position from the blade  50 . The vane assembly  54  includes one or more vanes  56  that prepare the airflow for the next set of blades. Additional vane assemblies could also be disposed within the portion  100 , including at a position upstream from the blade  50 . 
     The blade  50  includes a blade tip  58  that is positioned at a radially outermost portion of the blade  50 . In this exemplary embodiment, the blade tip  58  includes a knife edge  60  that extends toward a blade outer air seal (BOAS)  72 . The BOAS  72  establishes an outer radial flow path boundary of the core flow path C. The knife edge  60  and the BOAS  72  cooperate to limit airflow leakage around the blade tip  58 . 
     The BOAS  72  is disposed in an annulus radially between the casing  52  and the blade tip  58 . Although this particular embodiment is illustrated in a cross-sectional view, the BOAS  72  may form a full ring hoop assembly that circumscribes associated blades  50  of a stage of the portion  100 . 
     A seal member  62  is mounted radially inward from the casing  52  to the BOAS  72  to limit the amount of airflow AF to the annular cavity formed by the casing  52  and the BOAS  72 . A second seal member  64  can also be used, in conjunction with a flowpath member, to limit the amount of airflow leakage into the core flow path C. The second seal member  64  can mountably receive the BOAS  72 . The seal member  62  can also press the BOAS  72  axially against the adjacent vane assembly  54 , which forms a seal between the BOAS  72  and the vanes  56  to further limit cooling air leakage into the core flow path C. 
     In this exemplary embodiment, a dedicated cooling airflow, such as bleed airflow, is not communicated to cool the BOAS  72 . Instead, as is further discussed below, the BOAS  72  can include cooling features that increase a local heat transfer effect of the BOAS  72  without requiring a large flow pressure ratio. 
       FIG. 3  illustrates one exemplary embodiment of a BOAS  72  that may be incorporated into a gas turbine engine, such as a gas turbine engine  20 . The BOAS  72  of this exemplary embodiment is a full ring BOAS that can be circumferentially disposed about the engine centerline longitudinal axis A. The BOAS  72  can be formed as a single piece construction using a casting process or some other manufacturing technique. The BOAS  72  could also be segmented to include a plurality of BOAS segments within the scope of this disclosure. 
     The BOAS  72  includes a seal body  80  having a radially inner face  82  and a radially outer face  84 . Once positioned within the gas turbine engine  20 , the radially inner face  82  faces toward the blade tip  58  (i.e., the radially inner face  82  is positioned on the core flow path side) and the radially outer face  84  faces the casing  52  (i.e., the radially outer face  84  is positioned on a non-core flow path side). The radially inner face  82  and the radially outer face  84  axially extend between a leading edge portion  86  and a trailing edge portion  88 . 
     The leading edge portion  86  and the trailing edge portion  88  may include one or more attachment features  94  for sealing the BOAS  72  to the seal member  62  ( FIG. 2 ). In this exemplary embodiment, the leading edge portion  86  includes a hook  92  that receives the second seal member  64  to seal the BOAS  72  to the flowpath member. 
     The BOAS  72  can also include one or more cooling fins  96  disposed on the radially outer face  84  of the seal body  80 . In this exemplary embodiment, the BOAS  72  includes a plurality of circumferentially spaced cooling fins  96 . The cooling fins  96  can extend between a length L that extends between the leading edge portion  86  and the trailing edge portion  88 . In one exemplary embodiment, the cooling fins  96  extend across the entire length L between the leading edge portion  86  and the trailing edge portion  88 . 
     The cooling fins  96  can be cast integrally with the radially outer face  84  of the seal body  80 . In one exemplary embodiment, the BOAS  72  is made of a material having a relatively low coefficient of thermal expansion. Example materials include, but are not limited to, Mar-M-247, Hastaloy N, Hayes 242 and PWA 1456 (IN792+Hf). Other materials may also be utilized within the scope of this disclosure. 
       FIG. 4  illustrates a portion of the BOAS  72  of  FIG. 3 . A seal  98  can be secured to the radially inner face  82  of the seal body  80 . The seal  98  can be brazed to the radially inner face  82 , or could be attached using other known attachment techniques. In one example, the seal  98  is a honeycomb seal that interacts with a blade tip  58  of a blade  50  (See  FIG. 2 ) to reduce airflow leakage around the blade tip  58 . 
     A thermal barrier coating  102  can also be applied to at least a portion of the radially inner face  82  and/or the seal  98 . In this exemplary embodiment, the thermal barrier coating  102  is applied to the radially inner face  82  between the leading edge portion  86  and the trailing edge portion  88 . The thermal barrier coating  102  could also partially or completely fill the seal  98  of the BOAS  72 . The thermal barrier coating  102  may also be deposited on any flow path connected portion of the BOAS  72  to protect the underlying substrate of the BOAS  72  from exposure to hot gas, reducing thermal fatigue and to enable higher operating conditions. A suitable low conductivity thermal barrier coating  102  can be used to increase the effectiveness of the cooling fins  92  by reducing the heat transfer from the core flow path C to the airflow AF. 
     The cooling fins  96  include an outer surface  91 . The outer surface  91  can include a stepped portion  93  such that each cooling fin  96  includes a varying height across its length L relative to the radially outer face  84  of the BOAS  72 . For example, as illustrated in this embodiment, the cooling fins  96  include a first height H 1  adjacent to the leading edge portion  86  and include a second height H 2  that is different than the first height H 1  adjacent to the trailing edge portion  88 . In one embodiment, the second height H 2  is smaller than the first height H 1 . 
     Airflow AF is provided to the engine static structure  33  through the seal member  62  and is communicated into the passage created between the casing  52  and the BOAS  72  to prevent hot combustion gases from the core flow path C from contacting the casing  52 . The airflow AF can be communicated across the length L of each cooling fin  96  to cool the BOAS  72  without requiring additional flow, or a dedicated source of cooling air. The cooling fins  96  increase the surface area of the BOAS  72 , thereby increasing the local heat transfer effect of the BOAS  72  without requiring a large flow pressure ratio. 
     Referring to the embodiment depicted by  FIG. 5 , the BOAS  72  can also include a plurality of cooling fins  96  that embody different lengths. In one exemplary embodiment, a first portion  96 A of the plurality of cooling fins  96  can include a first length L 1 , while a second portion  96 B of the plurality of cooling fins  96  includes a second length L 2  that is greater than the first length L 1 . The first portion  96 A of the plurality of cooling fins  96  can be machined down to the length L 1  to provide clearance for mounting the BOAS to the casing  52 . The actual dimensions of the lengths L 1  and L 2  may be design dependent. 
       FIG. 6  illustrates additional features that may be incorporated into the BOAS  72 . In this exemplary embodiment, a portion of the cooling fins  96  can extend at a non-perpendicular angle α 1  relative to the radially outer face  84 , while another portion of the cooling fins  96  may extend at a perpendicular angle α 2  relative to the radially outer face  84 . The actual values of the angles α 1  and α 2  may be design dependent. 
     Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
     The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would recognize that various modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.