Abstract:
A wall assembly is provided for a gas turbine engine. This wall assembly includes a support shell with a contoured region; and a multiple of liner panels mounted to the support shell. At least one of the multiple of liner panels includes an end rail. The contoured region is deformable to selectively contact at least a portion of the end rail. A method of assembling a wall assembly within a gas turbine engine is also provided. This method includes locating a stud that extends from a cold side of a liner panel through a support shell; and attaching a fastener onto the stud to at least partially close a gap defined between the panel and shell.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. patent application Ser. No. 61/912,878 filed Dec. 6, 2013, which is hereby incorporated herein by reference in its entirety. 
     
    
     BACKGROUND 
       [0002]    The present disclosure relates to a gas turbine engine and, more particularly, to a hot section component therefor. 
         [0003]    Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. 
         [0004]    Among the engine components, relatively high temperatures are observed in the combustor section such that cooling is required to meet service life requirements. The combustor section typically includes an outer shell lined with heat shields often referred to as floatwall panels. In certain arrangements, dilution holes in the floatwall panel communicate with respective dilution holes in the outer shell to direct cooling air for dilution of the combustion gases. In addition to the dilution holes, the outer shell may also have relatively smaller air impingement holes to direct cooling air between the floatwall panels and the outer shell to cool the cold side of the floatwall panels. This cooling air exits effusion holes through of the floatwall panels to form a film on a hot side of the floatwall panels as a barrier against the hot combustion gases. 
         [0005]    With lower emission requirements and higher combustor temperatures, the amount of cooling is reduced while the effectiveness thereof is increased. A challenge related to this concept is that there is a higher pressure drop across the combustor panel. This higher pressure drop may result in increased sensitivity to leakage. 
       SUMMARY 
       [0006]    A wall assembly for a gas turbine engine, according to one disclosed non-limiting embodiment of the present disclosure, includes a support shell with a contoured region. The wall assembly also includes a multiple of liner panels mounted to the support shell. At least one of the multiple of liner panels includes an end rail. The contoured region is deformable to selectively contact at least a portion of the end rail. 
         [0007]    In a further embodiment of the present disclosure, the portion of the end rail and the contoured region of the support shell define a gap when the contoured region is in a first position. 
         [0008]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the portion of the end rail and the contoured region of the support shell form a seal when the contoured region is in a second position. 
         [0009]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the contoured region is centered with respect to an intermediate circumferential rail of the at least one of the multiple of liner panels. 
         [0010]    In a further embodiment of any of the foregoing embodiments of the present disclosure, a perimeter rail is included that defines a first height from a cold side of the at least one liner panel. The intermediate rail defines a second height from the cold side. The second height is less than the first height. 
         [0011]    In a further embodiment of any of the foregoing embodiments of the present disclosure, a multiple of studs are included that extend from each of the multiple of liner panels. 
         [0012]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the multiple of studs extend from the at least one of the multiple of liner panels generally along the intermediate circumferential rail. 
         [0013]    A wall assembly for a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes a support shell with a contoured region; a multiple of liner panels mounted to the support shell; and a multiple of fasteners mounted to the at least one of the multiple of liner panels. At least one of the multiple of liner panels includes a first and second end rail. The multiple of fasteners are operable to deform the contoured region to seal the support shell to the first and second end rail. 
         [0014]    In a further embodiment of any of the foregoing embodiments of the present disclosure, a multiple of studs are included that extend from the at least one of the multiple of liner panels through the support shell. Each of the multiple of fasteners respectively received onto one of the multiple of studs. 
         [0015]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the contoured region is centered with respect to an intermediate circumferential rail of the at least one of the multiple of liner panels. 
         [0016]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the liner panel is mounted within a combustor of the gas turbine engine. 
         [0017]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the first and the second end rails form a portion of a perimeter rail. 
         [0018]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the perimeter rail surrounds an intermediate rail. 
         [0019]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the perimeter rail includes a forward circumferential rail and an aft circumferential rail connected to the first and second end rails. The first and second end rails each include an engagement area in contact with the support shell when the contoured region of the support shell is in a first position and when the contoured region is in a second position sealed with the first and second end rail. 
         [0020]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the intermediate rail is generally parallel to the forward circumferential rail and the aft circumferential rail. 
         [0021]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the intermediate rail extends less than the forward circumferential rail and the aft circumferential rail by about 0.005-0.020 inches (0.1-0.5 mm) 
         [0022]    A method of assembling a wall assembly within a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure, includes locating a stud that extends from a cold side of a liner panel through a support shell; and attaching a fastener onto the stud to at least partially close a gap defined between the panel and shell. 
         [0023]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes at least partially closing the gap includes elastically deforming the support shell toward the liner panel. 
         [0024]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the support shell includes a contoured region and the panel includes a rail. The rail and the contoured region define a gap when the contoured region is in a first position and contact one another when the contoured region is in a second position. 
         [0025]    In a further embodiment of any of the foregoing embodiments of the present disclosure, the method includes defining the gap adjacent to an intermediate rail. 
         [0026]    The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
           [0028]      FIG. 1  is a schematic cross-section of an example gas turbine engine architecture; 
           [0029]      FIG. 2  is a schematic cross-section of another example gas turbine engine architecture; 
           [0030]      FIG. 3  is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment that may be used with the example gas turbine engine architectures shown in  FIGS. 1 and 2 ; 
           [0031]      FIG. 4  is an exploded partial sectional view of a portion of a combustor wall assembly; 
           [0032]      FIG. 5  is a perspective view of a portion of a liner panel array; 
           [0033]      FIG. 6  is an expanded longitudinal sectional view of a forward outer portion of a combustor wall assembly illustrating a contoured region of a support shell according to one disclosed non-limiting embodiment; 
           [0034]      FIG. 7  is a perspective partial view of one forward outer liner panel mounted to a support shell; 
           [0035]      FIG. 8  is a perspective view of a forward outer liner panel of a combustor wall assembly according to one disclosed non-limiting embodiment; 
           [0036]      FIG. 9  is a perspective view of a forward outer liner panel of a combustor wall assembly according to another disclosed non-limiting embodiment; 
           [0037]      FIG. 10  is an expanded longitudinal sectional view of a forward outer portion of a combustor wall assembly illustrating a gap according to one disclosed non-limiting embodiment with a contoured region of the shell in a first relaxed position; 
           [0038]      FIG. 11  is an expanded longitudinal sectional view of a forward outer portion of a combustor wall assembly illustrating the contoured region of the shell in a second deformed position, closing the gap of  FIG. 10 ; 
           [0039]      FIG. 12  is an expanded longitudinal sectional view of a forward outer portion of a combustor wall assembly illustrating the gap between the end rail and the contoured region of the shell in the first relaxed position; and 
           [0040]      FIG. 13  is an expanded longitudinal sectional view of a forward outer portion of a combustor wall assembly illustrating the contoured region of the shell in a second deformed position, closing the gap. 
       
    
    
     DETAILED DESCRIPTION 
       [0041]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbo fan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Referring to  FIG. 2 , alternative engine architectures  200  might include an augmentor section  12 , an exhaust duct section  14  and a nozzle section  16  in addition to the fan section  22 ′, compressor section  24 ′, combustor section  26 ′ and turbine section  28 ′ among other systems or features. Referring again to  FIG. 1 , the fan section  22  drives air along a bypass flowpath and into the compressor section  24 . The compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26 , which then expands and directs the air through the turbine section  28 . Although depicted as a turbofan in the disclosed non-limiting embodiment, 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 such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an Intermediate Pressure Compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an Intermediate Pressure Turbine (“IPT”) between a High Pressure Turbine (“HPT”) and a Low Pressure Turbine (“LPT”). 
         [0042]    The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing structures  38  or systems. The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor (“LPC”)  44  and a low pressure turbine (“LPT”)  46 . The inner shaft  40  may drive the fan  42  directly or through a geared architecture  48  as illustrated in  FIG. 1  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
         [0043]    The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (“HPC”)  52  and a high pressure turbine (“HPT”)  54 . A combustor  56  is arranged between the HPC  52  and the HPT  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0044]    Core airflow is compressed by the LPC  44  then the HPC  52 , mixed with the fuel and burned in the combustor  56 , then expanded over the HPT  54  and the LPT  46 . The LPT  46  and HPT  54  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. The main engine shafts  40 ,  50  are supported at a plurality of points by the bearing structures  38  within the static structure  36 . 
         [0045]    In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the LPC  44  and LPT  46  and render increased pressure in a fewer number of stages. 
         [0046]    A pressure ratio associated with the LPT  46  is pressure measured prior to the inlet of the LPT  46  as related to the pressure at the outlet of the LPT  46  prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC  44 , and the LPT  46  has a pressure ratio that is greater than about five (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 disclosure is applicable to other gas turbine engines including direct drive turbofans. 
         [0047]    In one embodiment, a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition —typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
         [0048]    Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“Tram”/518.7) 0.5 . The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
         [0049]    With reference to  FIG. 3 , the combustor section  26  generally includes a combustor  56  with an outer combustor wall assembly  60 , an inner combustor wall assembly  62  and a diffuser case module  64 . The outer combustor wall assembly  60  and the inner combustor wall assembly  62  are spaced apart such that a combustion chamber  66  is defined therebetween. The combustion chamber  66  is generally annular in shape to surround the engine central longitudinal axis A. 
         [0050]    The outer combustor liner assembly  60  is spaced radially inward from an outer diffuser case  64 A of the diffuser case module  64  to define an outer annular plenum  76 . The inner combustor liner assembly  62  is spaced radially outward from an inner diffuser case  64 B of the diffuser case module  64  to define an inner annular plenum  78 . It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto. 
         [0051]    The combustor wall assemblies  60 ,  62  contain the combustion products for direction toward the turbine section  28 . Each combustor wall assembly  60 ,  62  generally includes a respective support shell  68 ,  70  which supports one or more liner panels  72 ,  74  mounted thereto that are arranged to form a liner array. The support shells  68 ,  70  may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell  68  and inner shell  70 . Each of the liner panels  72 ,  74  may be generally rectilinear with a circumferential arc. The liner panels  72 ,  74  may be manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material. In one disclosed non-limiting embodiment, the liner array includes a multiple of forward liner panels  72 A and a multiple of aft liner panels  72 B that are circumferentially staggered to line the outer shell  68 . A multiple of forward liner panels  74 A and a multiple of aft liner panels  74 B are circumferentially staggered to line the inner shell  70 . 
         [0052]    The combustor  56  further includes a forward assembly  80  immediately downstream of the compressor section  24  to receive compressed airflow therefrom. The forward assembly  80  generally includes a cowl  82 , a bulkhead assembly  84 , and a multiple of swirlers  90  (one shown). Each of the swirlers  90  is circumferentially aligned with one of a multiple of fuel nozzles  86  (one shown) and the respective hood ports  94  to project through the bulkhead assembly  84 . 
         [0053]    The bulkhead assembly  84  includes a bulkhead support shell  96  secured to the combustor walls  60 ,  62 , and a multiple of circumferentially distributed bulkhead liner panels  98  secured to the bulkhead support shell  96  around the swirler opening The bulkhead support shell  96  is generally annular and the multiple of circumferentially distributed bulkhead liner panels  98  are segmented, typically one to each fuel nozzle  86  and swirler  90 . 
         [0054]    The cowl  82  extends radially between, and is secured to, the forwardmost ends of the combustor walls  60 ,  62 . The cowl  82  includes a multiple of circumferentially distributed hood ports  94  that receive one of the respective multiple of fuel nozzles  86  and facilitates the direction of compressed air into the forward end of the combustion chamber  66  through a swirler opening  92 . Each fuel nozzle  86  may be secured to the diffuser case module  64  and project through one of the hood ports  94  and through the swirler opening  92  within the respective swirler  90 . 
         [0055]    The forward assembly  80  introduces core combustion air into the forward section of the combustion chamber  66  while the remainder enters the outer annular plenum  76  and the inner annular plenum  78 . The multiple of fuel nozzles  86  and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber  66 . 
         [0056]    Opposite the forward assembly  80 , the outer and inner support shells  68 ,  70  are mounted to a first row of Nozzle Guide Vanes (NGVs)  54 A in the HPT  54 . The NGVs  54 A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section  28  to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs  54 A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed. 
         [0057]    With reference to  FIG. 4 , a multiple of studs  100  extend from each of the liner panels  72 ,  74  so as to permit an array (see  FIG. 5 ) of the liner panels  72 ,  74  to be mounted to their respective support shells  68 ,  70  with fasteners  102  such as nuts. That is, the studs  100  project rigidly from the liner panels  72 ,  74  and through the respective support shells  68 ,  70  to receive the fasteners  102  at a threaded section thereof 
         [0058]    A multiple of cooling impingement passages  104  penetrate through the support shells  68 ,  70  to allow air from the respective annular plenums  76 ,  78  to enter cavities  106  formed in the combustor walls  60 ,  62  between the respective support shells.  68 ,  70  and liner panels  72 ,  74 . The cooling impingement passages  104  are generally normal to the surface of the liner panels  72 ,  74 . The air in the cavities  106  provide cold side impingement cooling of the liner panels  72 ,  74  that is generally defined herein as heat removal via internal convection. 
         [0059]    A multiple of effusion passages  108  penetrate through each of the liner panels  72 ,  74 . The geometry of the passages (e.g., diameter, shape, density, surface angle, incidence angle, etc.) as well as the location of the passages with respect to the high temperature main flow also contributes to effusion film cooling. The combination of impingement passages  104  and effusion passages  108  may be referred to as an Impingement Film Floatwall (IFF) assembly. 
         [0060]    The effusion passages  108  allow the air to pass from the cavities  106  defined in part by a cold side  110  of the liner panels  72 ,  74  to a hot side  112  of the liner panels  72 ,  74  and thereby facilitate the formation of a thin, cool, insulating blanket or film of cooling air along the hot side  112 . The effusion passages  108  are generally more numerous than the impingement passages  104  to promote the development of film cooling along the hot side  112  to sheath the liner panels  72 ,  74 . Film cooling as defined herein is the introduction of a relatively cooler air at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the region of the air injection as well as downstream thereof. 
         [0061]    A multiple of dilution passages  116  may penetrate through both the respective support shells  68 ,  70  and liner panels  72 ,  74  along a common axis. For example only, in a Rich-Quench-Lean (R-Q-L) type combustor, the dilution passages  116  are located downstream of the forward assembly  80  to quench the hot combustion gases within the combustion chamber  66  by direct supply of cooling air from the respective annular plenums  76 ,  78 . 
         [0062]    With reference to  FIG. 6 , in one disclosed non-limiting embodiment, each of the respective support shells  68 ,  70  are at least partially non-parallel with respect to the forward liner panels  72 A,  74 A to form a convergent passage  120  therebetween along a contoured region  124 . That is, the contoured region  124  is a radially displaced profile section of the respective support shells  68 ,  70 . Although the forward liner panels  72 A, and the respective forward portion of the outer support shell  68  of the outer wall assembly will be specifically described and illustrated in each of the disclosed non-limiting embodiments, it should be appreciated that the inner support shell  70  and associated forward liner panels  72 B of the inner wall assembly may alternatively or additionally benefit herefrom. In addition, various other wall assemblies within a gas turbine engine such as within the walls of the augmentor section  12 , the exhaust duct section  14  and the nozzle section  16  (see  FIG. 2 ) may alternatively or additionally benefit herefrom. That is, the contoured region  124  and interface therefor may alternatively or additionally be located within engine sections other than the combustor section  26  which utilize a support shell, liner panel type wall arrangement. The various contoured regions  124  of the support shell  68  form one or more convergent passages  120  for panel cooling air by varying the profile of the combustor shell adjacent to the respective liner panels. Various contours and configurations are possible to tailor the location of the effusion air exit, and optimize heat transfer, pressure loss, manufacturability, NOx reduction, etc. Beneficially, the countered regions do not require additional hardware over conventional float wall combustor panels to create the convergence and are readily produced with current manufacturing methods. 
         [0063]    In this disclosed non-limiting embodiment, the contoured region  124  of the support shell  68  defines a hyperbolic cosine (COSH) profile in longitudinal cross-section that extends away from the forward liner panels  72 A. That is, the forward liner panels  72 A are generally linear in longitudinal cross-section, while the contoured region  124  is non-linear in longitudinal cross-section. For perspective, in this disclosed non-limiting embodiment, each of the forward liner panels  72 A define an axial length of about 1.5 inches (38 mm) and each may extend over a circumferential arc of about forty (40) degrees (one shown in  FIGS. 7 and 8 ). 
         [0064]    The contoured region  124  is located adjacent to a row of studs  100 A and an intermediate circumferential rail  126  is located between a forward circumferential rail  128  and an aft circumferential rail  130 . Each of the studs  100 A may be at least partially surrounded by posts  132  to at least partially support and operate as stand-offs between the support shell  68  and the forward liner panels  72 A. 
         [0065]    With reference to  FIG. 7 , each of the forward liner panels  72 A, in one disclosed non-limiting embodiment, includes a single row of studs  100 A (five shown) that extend through respective stud apertures  134  in the support shell  68 . A center or “king” stud  100 Aa is received within a central circular stud aperture  134   a  while the remainder of the studs  100 Ab are received within elongated apertures  134   b  to facilitate operational thermal growth relative to the center or “king” stud  100 Aa (see  FIG. 7 ). 
         [0066]    With continued reference to  FIG. 6 , the contoured region  124  forms a cavity  106 A that converges toward the forward circumferential rail  128  and the aft circumferential rail  130 . The cavity  106 A is further subdivided by the intermediate circumferential rail  126  into a forward cavity  106 Aa and an aft cavity  106 Ab. The forward cavity  106 Aa and the aft cavity  106 Ab thereby accelerate and direct impingement airflow from impingement passages  104  on each respective side of the intermediate circumferential rail  126  toward forward effusion apertures  108   a  and aft effusion apertures  108   b.  The forward effusion apertures  108   a  and the aft effusion apertures  108   b  may define respective angles through the forward liner panels  72 A to direct effusion airflow generally forward and aft into the combustion chamber  66 . It should be appreciated that various contours and configurations are possible to tailor the location of the effusion air passages to optimize heat transfer, pressure loss, manufacturability, etc., without need for additional hardware between the respective support shell and the liner panels. 
         [0067]    With reference to  FIG. 8 , in one disclosed non-limiting embodiment, the multiple of studs  100 A extend generally along the intermediate rail  126 . That is, the studs  100 A are axially aligned with, and may at least partially form, the intermediate rail  126 . In another disclosed non-limiting embodiment, a forward row of studs  100 Ba extend from the cold side  110  on one side of the intermediate rail  126  and a second row of studs  100 Bb that extend from the cold side  110  on a side of the intermediate rail  126  opposite the forward row of studs  100 Ba (see  FIG. 9 ). 
         [0068]    End rails  136  circumferentially close-out each forward liner panels  72 A with respect to the support shell  68 . That is, the forward circumferential rail  128  and the aft circumferential rail  130  are located at relatively constant curvature axial interfaces while the end rails  136  extend across an axial length of the support shell  68  to complete a perimeter rail  138  that seals the periphery of each forward liner panels  72 A with respect to the respective support shell  68 . 
         [0069]    With reference to  FIG. 10 , the intermediate rail  126  extends from the cold side  110  with respect to the forward circumferential rail  128 , the aft circumferential rail  130  and the end rails  136 . The intermediate rail  126  may extend for a lesser distance from the cold side  110  to form an offset with respect to the contoured region  124 . The intermediate rail  126  extends from the cold side  110  and is offset from the support shell  68  when the contoured region is in a first position and contacts the support shell  68  when the contoured region is in a second deformed/deflected position. The intermediate rail  126  thereby forms a gap G (e.g., a preassembly gap) that will cause the forward liner panels  72 A and the respective support shell  68  to deflect toward each other as the fasteners  102  are tightened onto the studs  100  to close the gap G (see  FIG. 11 ). It should be appreciated that fasteners such as clips and mechanisms other than threads may alternatively or additionally be utilized. 
         [0070]    As the fasteners  102  are tightened onto the studs  100 , the attachment at least partially elastically deflects the support shell  68  adjacent to the intermediate rail  126  and produces a tight seal between the perimeter rail  138  and the support shell  68  to assure an effective seal therebetween (See  FIGS. 11 and 13 ). The elastic deformation of the support shell  68  to seal with the intermediate rail  126  reduces—or eliminates—leakage to facilitate formation of relatively large pressure drops across the liner panels  72 ,  74  and thereby increase cooling effectiveness. 
         [0071]    With reference to  FIG. 12 , each end rail  136  and the contoured region  124  define a respective rail profile  140  and support shell profile  142  contoured to form the gap G prior to engagement of the fasteners  102 . That is, the rail profile  140  of the end rails  136  and the support shell profile  142  of the contoured region  124  are of a slightly different profile such that the gap G is located in an area  150  generally axially adjacent to the intermediate rail  126  to form respective axial engagement areas  152 A,  152 B to provide a sliding engagement adjacent to the forward circumferential rail  128  and the aft circumferential rail  130  (See  FIGS. 10 and 11 ). That is, the support shell profile  142  slides along the respective axial engagement areas  152 A,  152 B as the fasteners  102  are engaged. In one example, the gap G is about 0.005-0.020 inches (0.1-0.5 mm). It should be further appreciated that the intermediate rail  126  may be of the same height from the cold side  110  as the forward circumferential rail  128  and the aft circumferential rail  130  or alternatively, of a lesser height to further facilitate formation of the gap G. 
         [0072]    The gap G is closed in response to deflection of the contoured region  124  as the fasteners  102  are tightened onto the studs  100  (see  FIG. 13 ). As the fasteners  102  are tightened down, they slightly deform the sheet metal support shell  68  against the rails. That is, tightening of the fasteners  102  closes the gap G such that the contoured region  124  closely follows the end rails  136  to circumferentially close-out and form an interference fit between each of the forward liner panels  72 A with respect to the associated support shell  68 . 
         [0073]    This interference fit ensures an effective seals that reduces leakage to facilitate formation of a relatively larger pressure drops across the liner panels  72 ,  74  and increase cooling effectiveness. 
         [0074]    The use of the terms “a” and “an” and “the” and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
         [0075]    Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention 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. 
         [0076]    It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
         [0077]    Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
         [0078]    The foregoing description is exemplary rather than defined by the features within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.