Patent Publication Number: US-9896956-B2

Title: Support assembly for a gas turbine engine

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with government support under Contract No. FA-8650-09-D-2923-0021 awarded by the United States Air Force. The Government has certain rights in this invention. 
    
    
     BACKGROUND 
     Gas turbine engines typically include a fan delivering air into a compressor. The air is compressed in the compressor and delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine blades, driving them to rotate. Turbine rotors, in turn, drive the compressor and fan rotors. 
     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. 
     Since the rotating blade and blade outer air seal may respond radially at different rates due to loads, the tip clearance may be reduced and the blade may rub on the blade outer air seal, which is undesirable. Therefore, there is a need to control the clearance between the blade and the blade outer air seal in order to increase the efficiency of the gas turbine engine. 
     SUMMARY 
     In one exemplary embodiment, a support assembly for a gas turbine engine includes at least one inner support that extends about a circumferential axis and defines a cavity for receiving a control ring. At least one cover plate is attached to at least one inner support to enclose the cavity. At least one of the inner support and the cover plate includes a rail and the other of the inner support and the cover plate includes a groove for engaging the rail. 
     In a further embodiment of the above, a control ring is located in the cavity and extends about the circumferential axis. 
     In a further embodiment of any of the above, the groove is located in the cover plate and the rail is located on the inner support. 
     In a further embodiment of any of the above, the groove is located in the inner support and the rail is located on the cover plate. 
     In a further embodiment of any of the above, at least one of the groove and the rail are located on a hook portion of the at least one cover plate. 
     In a further embodiment of any of the above, the inner support includes a recess for accepting the hook portion. 
     In a further embodiment of any of the above, at least one of the groove and the rail is circumferentially offset from the recess. 
     In a further embodiment of any of the above, the groove and the rail are axially aligned with the recess. 
     In a further embodiment of any of the above, a blade outer air seal is attached to the inner support. 
     In a further embodiment of any of the above, at least one cover plate and the at least one inner support are made of materials with similar metal properties. 
     In a further embodiment of any of the above, at least one cover plate and the at least one inner support are made of the same material. 
     In another exemplary embodiment, a gas turbine engine includes at least one inner support that extends about a circumferential axis and defines a cavity for receiving a control ring. At least one cover plate is attached to at least one inner support to enclose the cavity. At least one of the inner support and the cover plate includes a rail and the other of the inner support and the cover plate includes a groove for engaging the rail. A blade outer air seal is attached to the inner support. 
     In a further embodiment of any of the above, a control ring is located in the cavity and extends about the circumferential axis. 
     In a further embodiment of any of the above, at least one of the groove and the rail are located on a hook portion of the at least one cover plate. 
     In a further embodiment of any of the above, the inner support includes a recess for accepting the hook portion. 
     In a further embodiment of any of the above, at least one of the groove and the rail is circumferentially offset from the recess. 
     In a further embodiment of any of the above, the groove and the rail are axially aligned with the recess. 
     In a further embodiment of any of the above, at least one cover plate and at least one inner support are made of materials with similar metal properties. 
     In a further embodiment of any of the above, at least one cover plate and the at least one inner support are made of the same material. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine. 
         FIG. 2  is a cross-sectional view of a turbine section of the example gas turbine engine of  FIG. 1 . 
         FIG. 3  is a cross-sectional view of an example support assembly for a blade outer air seal. 
         FIG. 4  is a perspective view of a portion of the support assembly of  FIG. 3 . 
     
    
    
     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 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  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 fan  42 , 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 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 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  is 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 combustor section  26  or even aft of turbine section  28 , and fan section  22  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. 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. 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). 
     The example gas turbine engine includes fan  42  that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment low pressure turbine  46  includes about three (3) turbine rotors. A ratio between number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate fan section  22  and therefore the relationship between the number of turbine rotors  34  in low pressure turbine  46  and number of blades  42  in fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Although the gas turbine engine  20  shown is a high bypass gas turbine engine, other types of gas turbine engines could be used, such as a turbojet engine. 
       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 with a first rotor assembly  60 . In another example, the high pressure turbine  54  could include a two or more stages high pressure turbine section. 
     The first rotor assembly  60  includes a first array of rotor blades  62  circumferentially spaced around a first disk  64 . Each of the first array of rotor blades  62  includes 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  68  of the first disk  64 . The first airfoil  80  extends radially outward toward a first blade outer air seal (BOAS) assembly  84 . The BOAS  84  is supported by a support assembly  100 . 
     The first array of rotor blades  62  are disposed in the core flow path 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 . 
     An array of vanes  90  are located axially upstream of the first array of rotor blades  62 . Each of the array of vanes  90  include at least one airfoil  92  that extend between a respective vane inner platform  94  and an vane outer platform  96 . In another example, each of the array of vanes  90  include at least two airfoils  92  forming a vane double. The vane outer platform  96  of the vane  90  may at least partially engage the BOAS  84 . 
     As shown in  FIGS. 2 and 3 , the support assembly  100  includes an outer support  102 , an inner support  104 , a control ring  106 , and a cover plate  108 . The outer support  102  forms a complete unitary hoop and includes an axially extending flange  110  and a radially extending flange  112 . The axially extending flange  110  engages a case or a portion of the engine static structure  36  when installed in the gas turbine engine  20 . The radially extending portion of the outer support  102  extends radially inward from the axially extending flange  110 . In this disclosure, radially or radially extending is in relation to the engine axis A of the gas turbine engine  20  unless stated otherwise. 
     The inner support  104  includes a C-shaped cross section with an opening of the C-shaped cross section facing an axially upstream or forward direction. The C-shaped cross section is formed by a radially inner flange  114  connected to a radially outer flange  116  by a radially extending flange  118 . The radially extending flange  118  includes an axial surface  120  that contacts or is in close proximity to an axial surface  122  on the radially extending flange  112  on the outer support  102  to prevent the inner support  104  from moving axially downstream past the radially extending flange  112 . 
     The radially outer flange  116  is spaced radially inward from the axially extending flange  110  on the outer support  102  such that a clearance between the axially extending flange  110  and the radially outer flange  116  is maintained during operation of the gas turbine engine  20 . By maintaining the clearance between the axially extending flange  110  and the radially outer flange  116 , the inner support  104  is allowed to grow radially outward when exposed to elevated operating temperatures during operation of the gas turbine engine  20  without transferring a load to the outer support  102 . 
     In the illustrated example, the radially inner flange  114  includes attachment members  124  that extend radially inward from a radially inner surface of the radially inner flange  114  to support the BOAS  84  as shown in  FIGS. 1 and 2 . Although the attachment members  124  are shown as a pair of hooks with distal ends pointing axially downstream in the illustrated example, the attachment members  124  could include hooks pointing in opposite directions or more than or less than two hooks. 
     In the illustrated example, the cover plate  108  is attached to an axially forward end of the inner support  104  to from a cavity  126  that surrounds the control ring  106 . Both the inner support  104  and the cover plate  108  are made of corresponding segments that fit together to form a circumferential ring. 
     In one example, the cover plate  108  and the inner support  104  are made of the same material or materials with similar metal properties. By making the cover plate  108  and the inner support  104  of the same material, the thermal growth of the cover plate  108  will closely match the thermal growth of the inner support  104  to ensure that the axial ends of the inner support  104  grow at a similar rate in the radial direction. 
     As shown in  FIGS. 2-4 , the cover plate  108  and the inner support  104  are attached to each other with a first retention member  130  and a second retention member  132 . In the illustrated example, the first retention member  130  includes a bayonet attachment portion  133  and a rail portion  135  on a radially outer edge of the cover plate  108  and the second retention member  132  includes a tab  134  on a radially inner edge of the cover plate  108 . The tab  134  extends in an axially downstream direction. The bayonet attachment portion  133  includes a hook portion  136  having a radially extending portion  136   a  that is axially offset from a body portion  138  of the cover plate  108 . The radially outer flange  116  of the inner support  104  includes a recess  140  for accepting the hook portion  136  and a groove  142  at least partially axially aligned with the recessed  140  and circumferentially offset such that the cover plate  108  can be rotated in a circumferential direction to move the hook portion  136  from the recessed  140  into the groove  142 . The radially extending portion  136   a  of the hook portion  136  engages axial faces of the groove  142  and an axially extending portion  136   b  of the hook portion  136  engages a radially outer surface of the radially outer flange  116 . 
     In the illustrated example, the rail portion  135  includes a rail  152  on a radially outer surface of the radially outer flange  116  of the inner support  104  and a groove  154  on the axially extending portion  136   b  of the hook portion  136  of the cover plate  108 . The rail  152  is circumferentially spaced from the recess  140  in the inner support  104 . This allows the hook portion  136  to slide over the outer flange  116  on the inner support  104  without interference from the rail  152 . The groove  154  in the hook portion  136  will then axially align with the rail  152  on the inner support  104  to allow the cover plate  108  to rotate relative to the inner support  104 . 
     In another example of the rail portion  135 , the rail  152  is located on the hook portion  136  and the groove  154  is located in the outer flange  116  of the inner support  104 . In yet a further example, the rail portion  135  could be located adjacent the inner flange  114  of the inner support  104  instead of the outer flange  116  or a first rail portion  135  could be located adjacent the inner flange  114  and a second rail portion  135  could be located adjacent the outer flange  116 . 
     The rail portion  135  allows for the cover plate  108  to be axially aligned with the inner support  104  and limits relative axial movement between the cover plate  108  and the inner support  104 . The rail portion  135  also improves installation by maintaining alignment of the cover plate  108  relative to the inner support  104 . 
     The tab  134 , which forms the second retention member  132 , is located on a radially inner edge of the cover plate  108 . The tab  134  engages a radially inner surface of the radially inner flange  114  on the inner support  104  such that the bayonet attachment portion  133  and the tab  134  surround the inner support  104 . 
     Opposing ends of the cover plate  108 , which are circumferentially spaced from the first retention member  130  and the second retention member  132 , fit within the inner support  104 . As shown in  FIGS. 3 and 4 , the opposing ends of the cover plate  108  contact a radially inner surface of the radially outer flange  116  and a radially inner edge of the cover plate  108  contacts a radially outer surface of the radially inner flange  114 . 
     During assembly of the support assembly  100 , the plurality of inner supports  104  are arranged in a circumferential ring surrounding the control ring  106  with the control ring  106  located in the cavity  126 . Each of the corresponding plurality of cover plates  108  is placed on the inner support  104  such that the hook portion  136  on each of the plurality of cover plates  108  is located within the corresponding recess  140  in each of the plurality of inner supports  104 . 
     When the plurality of cover plates  108  are located on the inner supports  104 , the plurality of cover plates  108  are rotated in unison such that the hook portion  136  with the groove  154  on each of the plurality of cover plates  108  moves into the corresponding grooves  142  on each of the inner supports  104  while the rail  152  moves into the groove  154 . When each of the plurality of cover plates  108  is initially placed in the grooves  142  of the inner support  104 , one of the circumferential ends of each of the plurality of cover plates  108  will overlap an adjacent inner support  104 . As the plurality of cover plates  108  rotate, each of the plurality of cover plates  108  will circumferentially align with a corresponding one of the inner supports  104 . The plurality of cover plates  108  are prevented from rotating further by a stop  144  on the inner support  104  that engages the tab  134 . 
     The inner supports  104 , the control ring  106 , and the plurality of cover plates  108  are then placed within the outer support  102  such that the axial surface  120  on the inner support  104  contacts or is in close proximity to the axial surface  122  on the outer support  102 . A plurality of cover plate tabs  150  extend from a radially inner surface of the axially extending flange  110  of the outer support  102  and engage an edge  156  on each of the hook portions  136  to prevent each of the cover plates  108  from rotating out of the groove  142  after being installed into the outer support  102 . 
     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.