Abstract:
This disclosure relates to a gas turbine engine including a blade outer air seal (BOAS) mounted radially outwardly of a blade. The engine further includes a BOAS support. A radial dimension of the BOAS support is selectively changeable in response to a flow of fluid through the BOAS support to adjust a radial position of the BOAS relative to the blade.

Description:
STATEMENT REGARDING GOVERNMENT SUPPORT 
       [0001]    This invention was made with government support under Contract No. FA8650-09-D-2923 0021 awarded by the United States Air Force. The government has certain rights in this invention. 
     
    
     BACKGROUND 
       [0002]    Gas turbine engines include turbine blades configured to rotate and extract energy from hot combustion gases that are communicated through the gas turbine engine. An outer casing of an engine static structure of the gas turbine engine may include one or more blade outer air seals (BOAS) that provide an outer radial flow path boundary for the hot combustion gases. 
       SUMMARY 
       [0003]    One exemplary embodiment of this disclosure relates to a gas turbine engine. The engine includes a blade outer air seal (BOAS) mounted radially outwardly of a blade, and a BOAS support. A radial dimension of the BOAS support is selectively changeable in response to a flow of fluid through the BOAS support to adjust a radial position of the BOAS relative to the blade. 
         [0004]    In a further embodiment of any of the foregoing, the BOAS support includes an inlet opening in a radially outer surface thereof, and an outlet opening in a radially outer surface thereof. 
         [0005]    In a further embodiment of any of the foregoing, the BOAS support includes a first chamber in communication with the inlet opening, and a second chamber in communication with the outlet opening. The first chamber in communication with the second chamber by way of an axial passageway. 
         [0006]    In a further embodiment of any of the foregoing, the engine includes an engine case. The engine case includes an inlet opening and an outlet opening aligned with a respective one of the inlet opening and outlet opening of the BOAS support. 
         [0007]    In a further embodiment of any of the foregoing, the engine includes a seal provided between a radially outer surface of the BOAS support and the engine case. 
         [0008]    In a further embodiment of any of the foregoing, the BOAS support is axially positioned between a first vane support and a second vane support. 
         [0009]    In a further embodiment of any of the foregoing, the engine includes a first seal provided between a fore surface of the BOAS support and the first vane support, and a second seal provided between an aft surface of the BOAS support and the second vane support. 
         [0010]    In a further embodiment of any of the foregoing, wherein the BOAS support includes at least three circumferentially spaced-apart flanges engaging a respective slot formed in the engine case to restrict circumferential movement of the BOAS support relative to the engine case. 
         [0011]    In a further embodiment of any of the foregoing, the flange includes a circumferential surface configured to contact a circumferential surface of the slot in at least one condition. 
         [0012]    In a further embodiment of any of the foregoing, the radial dimension of the BOAS support is changeable in response to a change in one of a temperature and a mass flow rate of the flow of fluid introduced into the BOAS support. 
         [0013]    Another exemplary embodiment of this disclosure relates to a blade outer air seal (BOAS) support include at least one flange extending from a radially outer surface of the BOAS support. The at least one flange is configured to radially support the BOAS support relative to an engine case, and to allow a radial dimension of the BOAS support to change. 
         [0014]    In a further embodiment of any of the foregoing, the at least one flange includes three circumferentially spaced-apart flanges. 
         [0015]    In a further embodiment of any of the foregoing, the BOAS support includes an inlet opening in a radially outer surface thereof, and an outlet opening in a radially outer surface thereof. 
         [0016]    In a further embodiment of any of the foregoing, the BOAS support includes a first chamber in communication with the inlet opening, and a second chamber in communication with the outlet opening. The first chamber is in communication with the second chamber by way of an axial passageway. 
         [0017]    In a further embodiment of any of the foregoing, the radial dimension of the BOAS support is changeable in response to a change in one of a temperature and a mass flow rate of a fluid introduced into the BOAS support. 
         [0018]    Another exemplary embodiment of this disclosure relates to a method for regulating tip clearance. The method includes regulating a clearance between a BOAS and a tip of a blade by changing a radial dimension of the BOAS support. The method further includes introducing fluid into a blade outer air seal (BOAS) support to change the radial dimension of the BOAS support. 
         [0019]    In a further embodiment of any of the foregoing, the method includes selectively changing the radial dimension of the BOAS support by changing one of a temperature and a mass flow rate of the fluid introduced into the BOAS support. 
         [0020]    In a further embodiment of any of the foregoing, the method includes increasing the radial dimension of the BOAS support in response to one of an increase in the temperature and a decrease in the mass flow rate of the fluid. 
         [0021]    In a further embodiment of any of the foregoing, the method includes decreasing the radial dimension of the BOAS support in response to one of a decrease in the temperature and an increase in the mass flow rate of the fluid. 
         [0022]    In a further embodiment of any of the foregoing, the fluid introduced into the BOAS support is a relatively lower temperature than a fluid introduced into the BOAS. 
         [0023]    The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]    The drawings can be briefly described as follows: 
           [0025]      FIG. 1  illustrates a schematic, cross-sectional view of an example gas turbine engine. 
           [0026]      FIG. 2  illustrates a cross-section of a portion of the gas turbine engine of  FIG. 1 . 
           [0027]      FIG. 3  is a close-up view of a portion of  FIG. 2 , and illustrates the detail of a BOAS support according to this disclosure. 
           [0028]      FIG. 4  is a view taken along line  4 - 4  in  FIG. 3 . 
       
    
    
     DETAILED DESCRIPTION 
       [0029]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes 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 or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws a core airflow C along a core flow path where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0030]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. The concepts disclosed herein can further be applied outside of gas turbine engines. 
         [0031]    The example 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. 
         [0032]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the 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 high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
         [0033]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0034]    The example low pressure turbine  46  has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0035]    A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0036]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow C and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0037]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0038]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0039]    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. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0040]    “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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
         [0041]    “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. 
         [0042]      FIG. 2  illustrates a portion  62  of the gas turbine engine  20  within a high pressure turbine  54 . 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 fan section  22 , the compressor section  24  and the low pressure turbine  46 . 
         [0043]    In this embodiment, a rotor disc  66  (only one shown, although multiple discs could be axially disposed within the portion  62 ) is mounted for rotation about the engine central longitudinal axis A. The portion  62  includes an array of rotating blades  68  (mounted to the rotor disc  66 ) positioned axially between arrays of vane assemblies  70 . The vane assemblies  70  each include a plurality of vanes  70 A,  70 B that are supported relative to an outer casing  69  of the engine static structure  36  ( FIG. 1 ) by way of first and second vane supports  72 ,  74 . 
         [0044]    Each blade  68  mounted to the rotor disc  66  includes a blade tip  68 T at radially outermost portion thereof. As referred to herein, the radial direction R is normal to the engine central longitudinal axis A. The rotor disc  66  is arranged such that the blade tip  68 T is located adjacent a blade outer air seal (BOAS) assembly  76 . The BOAS assembly  76  may find beneficial use in many industries including aerospace, industrial, electricity generation, naval propulsion, pumps for gas and oil transmission, aircraft propulsion, vehicle engines and stationary power plants. 
         [0045]    The BOAS assembly  76  is disposed in an annulus radially between the outer casing  69  and the blade tip  68 T. The BOAS assembly  76  in this example includes a BOAS support  78  and a multitude of BOAS segments  80  (only one shown in  FIG. 2 ). The BOAS segments  80  may be arranged to form a full ring hoop assembly that circumferentially surrounds the associated blade  68 . The BOAS support  78  is mounted radially inward from the outer casing  69 , and includes forward and aft flanges  78 A,  78 B that receive forward and aft flanges  80 A,  80 B of the BOAS segments  80 . In one example, a single BOAS support  78  provides a full hoop (or, full ring) around the engine central longitudinal axis A. Providing the BOAS support  78  as a full hoop restricts radial inward movement of the BOAS support  78 . The detail of the BOAS support  78 , including the manner in which the BOAS support  78  is mounted relative to the portion  62  of the engine  20  is illustrated in detail in  FIG. 3 . 
         [0046]    As illustrated in  FIG. 3 , the BOAS support  78  is provided axially between a first, forward vane support  72 , and a second, aft vane support  74 . The BOAS support  78  is provided radially inward of an outer casing  69 . It should be understood that the BOAS support  78  may be formed by casting, or by an additive manufacturing process. Other manufacturing techniques can be used to form the BOAS support  78 , however. 
         [0047]    In this example, a first seal  82  is provided between the first vane support  72  and the BOAS support  78 , and a second seal  84  is provided axially between the second vane support  74  and the BOAS support  78 . While only one of each of the first and second seals  82 ,  84  is illustrated, additional, redundant seals may be provided if desired. 
         [0048]    The BOAS support  78  is radially supported relative to the outer casing  69  in this example by way of a plurality of flanges  86 A- 86 C. The detail of the flanges  86 A- 86 C is illustrated in  FIG. 4 , which is a view taken along line  4 - 4  in  FIG. 3 . 
         [0049]    As shown in  FIG. 4 , a plurality of circumferentially spaced-apart flanges  86 A- 86 C extend radially outward, in a direction substantially normal to the engine central longitudinal axis A, from a radially outer surface  94  of the BOAS support  78 . The outer casing  69  similarly includes a plurality of radially extending slots  88 A- 88 C. The flanges  86 A- 86 C are received in a corresponding one of the slots  88 A- 88 C such that, in at least one condition, a radially outer surface  86 R of the each of the flanges  86 A- 86 C is spaced by a distance D 1  from an radially outer surface  88 R of the slot  88 A- 88 C to provide room to accommodate any relative expansion between the BOAS support  78  and the outer casing  69 . 
         [0050]    The outer flanges  86 A,  86 C also serve to circumferentially restrict movement of the BOAS support  78  relative to the outer casing  69 , by way of the outer circumferential faces  90  thereof. The outer circumferential faces  90  of the outer flanges  86 A,  86 C engage corresponding outer circumferential faces  92  of the outer slots  88 A,  88 C to restrict circumferential movement (e.g., such that the BOAS support  78  is provided substantially concentric with the outer casing  69 ). While three flanges  86 A- 86 C and three slots  88 A- 88 C are illustrated, it should be understood that any number of flanges and slots can be included. For instance, some examples may only include the outer flanges  86 A,  86 C while omitting the middle flange  86 B. Other examples may include greater numbers of flanges and slots, on the order of six or eight. 
         [0051]    With reference back to  FIG. 3 , the BOAS support  78  generally includes a radially outer surface  94  adjacent the flange  86 , a radially inner surface  96  adjacent the forward and aft attachment flanges  78 A,  78 B, a fore surface  98 , and an aft surface  100 . The various surfaces  94 ,  96 ,  98 ,  100  of the BOAS support  78  may optionally be coated with a thermal barrier coating (TBC) to insulate the BOAS support from the relatively hot gases adjacent the BOAS  80 . 
         [0052]    The fore surface  98  contacts the first seal  82 , and the aft surface  100  contacts the second seal  84 . In this example, the first and second seals  82 ,  84  each include at least one trough  82 T,  84 T facing radially inward to substantially prevent the relatively high temperature and high pressure BOAS cooling flow from interacting with the BOAS support  78 . Here, the first and second seals  82 ,  84  each include three troughs  82 T,  84 T. It should be understood that different types of seals, including seals with different numbers of troughs, come within the scope of this disclosure. 
         [0053]    The BOAS support  78  further includes a first chamber  102  and a second chamber  104 . The first and second chambers  102 ,  104  are in communication with one another via an axial passageway  106 . The first and second chambers  102 ,  104  may include turbulators, such as trip strips or pedestals, to increase heat transfer. The radially outer surface  94  of the BOAS support  78  includes an inlet opening  108  in communication with the first chamber  102 , and an exit opening  110  in communication with the second chamber  104 . 
         [0054]    A third seal  112  is provided axially between the inlet opening  108  and the exit opening  110 , and further extends in a radial direction R from a seat  112 S formed in the radially outer surface  94  to the outer casing  69 . In this example, the seal  112  is a piston ring seal, although different seals come within the scope of this disclosure. The outer casing  69  further includes inlet and outlet openings  114 ,  116  axially and circumferentially aligned with the inlet and outlet openings  108 ,  110  of the BOAS support  78 . 
         [0055]    During operation, a cooling flow S enters the inlet  114  in the outer casing  69 , and flows into the inlet  108  in the BOAS support  78 . The cooling flow S travels circumferentially through the first chamber  102 , and is next directed from the first chamber  102  to the second chamber  104  via the axial passageway  106 . The cooling flow S then travels circumferentially through the second chamber  104 , and finally passes out the exit openings  110  and  116 . The seal  112  prevents intermixing between the cooling flow S entering the BOAS support  78  and that exiting the BOAS support  78 . 
         [0056]    In one example, the cooling flow S is relatively low pressure air. This low pressure air reduces the work required to generate such air. This air may further be relatively cool (e.g., low temperature). In the example, the cooling flow S is relatively low pressure and low temperature compared to a flow of fluid used to cool the BOAS  80 . In the event of a seal failure (e.g., a failure of one of the seals  82 ,  84 ), the BOAS support  78  will be cooled by the fluid intended to cool the BOAS  80 , which still provides adequate cooling. In one example, the engine  20  includes a dedicated supply of fluid providing the cooling flow S. For example, the cooling flow S may be sourced from a fan bypass air. In this example, a valve may be introduced into the engine  20  to selectively tap air from a selected location for use in cooling the BOAS support  78 . Downstream of the BOAS support  78 , the cooling flow S is routed to an even lower pressure location, such as a third fluid flow stream. In another example, cooling flow S is tapped from an existing cooling system for another component of the engine  20 , such as that of a downstream vane or nozzle. 
         [0057]    During engine operation, it is extremely important to regulate the clearance of the blade tips  68 T relative to the BOAS segments  80 . For example, if the clearance between the blade tips  68 T and the BOAS segments  80  is too large, the engine  20  will operate inefficiently. On the other hand, if the clearance is too small, there may be excessive rubbing between the blades  68  and the BOAS segments  80  which can increase wear on the engine components. The tip clearance is monitored in some examples by direct measuring (e.g., by way of sensors), and in other examples by monitoring the efficiency of the engine  20 . 
         [0058]    In either case, the temperature or mass flow rate of the cooling flow S may be selectively changed (e.g., by selecting an alternate source for the cooling flow S, or by selectively mixing sources of fluid to provide the cooling flow S), to either radially expand (e.g., radially outward movement) or contract (e.g., radially inward movement) the BOAS support  78 , thus changing the radial dimension of the BOAS support  78 . 
         [0059]    For example, if the clearance is too small, a relatively higher temperature cooling flow S (or, a reduced mass flow rate) would be introduced into the BOAS support  78  to expand the BOAS support  78 , and increase the radial dimension thereof, essentially moving the BOAS support  78  away from the central longitudinal axis A. On the other hand, if the clearance is too large, a relatively low temperature cooling flow S (or, an increased mass flow rate) would be used to contract the BOAS support  78  and reduce the radial dimension thereof (moving the BOAS support  78  toward the central longitudinal axis A). The change in the radial dimension of the BOAS support  78  will correspond to a change in the radial clearance between the BOAS segments  80  and the blade tips  68 T. 
         [0060]    In addition to requiring a reduced work to provide the cooling flow S, this disclosure has the added benefit of having relatively few moving parts and being lightweight relative to other types of clearance control systems. Accordingly, this disclosure provides an effective, and inexpensive tip clearance regulation system. 
         [0061]    Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0062]    One of ordinary skill in this art would understand that the above-described embodiments are exemplary and non-limiting. That is, modifications of this disclosure would come within the scope of the claims. Accordingly, the following claims should be studied to determine their true scope and content.