Abstract:
A heat shield for a combustor of a gas turbine engine includes a first edge with a first set of cantilevered members and a second edge with a second set of cantilevered members.

Description:
[0001]    This application claims priority to U.S. Patent Appln. No. 61/762,367 filed Feb. 8, 2013. 
     
    
     BACKGROUND 
       [0002]    The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor. 
         [0003]    Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases. 
         [0004]    As engine requirements increase for improved thrust specific fuel consumption (TSFC), compressor discharge pressure and temperature along with combustor exit temperatures (CET) may also increase. As a result, current combustor configurations emissions, such as NOx, CO, unburned hydrocarbons (UHC), and smoke, may increase relative to exceedingly stringent emissions standards. 
       SUMMARY 
       [0005]    A heat shield for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a first edge with a first set of cantilevered members and a second edge with a second set of cantilevered members interleaved with the first set of cantilevered members. 
         [0006]    A further embodiment of any of the foregoing embodiments of the present disclosure includes interleaved cantilevered members between the first heat shield and the second heat shield. 
         [0007]    A further embodiment of any of the foregoing embodiments of the present disclosure includes an exit splitter that extends from the heat shield. 
         [0008]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the exit splitter is zigzag in shape. 
         [0009]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, further comprising a film hole located in a valley on each side of the exit splitter. 
         [0010]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, further comprising a plurality of studs which extend from the heat shield, the stud includes a frustro-conical section. 
         [0011]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the heat shield includes a multiple of pin fins. 
         [0012]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the multiple of pin fins are diamond-shaped. 
         [0013]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the heat shield includes a multiple of hemi-spherical nubbins. 
         [0014]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the multiple of hemi-spherical nubbins decrease in diameter toward an exit splitter. 
         [0015]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein a center of the sphere of each of the multiple of hemi-spherical nubbins are further displaced from an inner surface of the heat shield toward an exit splitter. 
         [0016]    A liner assembly for a combustor of a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a first heat shield and a second heat shield interleaved with the first heat shield. 
         [0017]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the second heat shield is interleaved with the first heat shield at interleaved cantilevered members between the first heat shield and the second heat shield. 
         [0018]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, further comprising a plurality of studs which extend from the first heat shield and are received through the first support shell, the plurality of studs include a frustro-conical section. 
         [0019]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, further comprising a nut which threads to the stud to force the first support shell onto the first heat shield. 
         [0020]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the nut threads to the stud to drive together the interleaved cantilevered members between the first heat shield and the second heat shield. 
         [0021]    A further embodiment of any of the foregoing embodiments of the present disclosure includes a first support shell with a non-planar profile which faces the first heat shield and a second support shell with a non-planar profile which faces the second heat shield 
         [0022]    A method of mounting a liner assembly of a combustor for a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes interleaving a first heat shield with a second heat shield. 
         [0023]    A further embodiment of any of the foregoing embodiments of the present disclosure includes interleaving a first set of cantilevered members that extend from the first heat shield with a second set of cantilevered members that extend from the second heat shield. 
         [0024]    A further embodiment of any of the foregoing embodiments of the present disclosure includes threading a nut to a stud that extends from the first heat shield to drive a support shell onto the first heat shield. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0025]    Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
           [0026]      FIG. 1  is a schematic cross-section of a gas turbine engine; 
           [0027]      FIG. 2  is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0028]      FIG. 3  is an expanded partial perspective longitudinal schematic view of a combustor section according to one non-limiting embodiment that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0029]      FIG. 4  is an exploded view of a liner assembly of the combustor; 
           [0030]      FIG. 5  is an expanded circumferentially partial perspective view of the combustor section associates with one pre-swirler; 
           [0031]      FIG. 6  is an expanded lateral sectional view of a liner assembly according to one non-limiting embodiment; 
           [0032]      FIG. 7  is an expanded lateral sectional view of the liner assembly of  FIG. 6  with a relationship for a non-planar profile that faces an inner surface of a heat shield of the liner assembly; 
           [0033]      FIG. 8  is an expanded plan view of a heat shield of a liner assembly according to one non-limiting embodiment; 
           [0034]      FIG. 9  is an expanded plan view of a heat shield of a liner assembly according to another non-limiting embodiment; 
           [0035]      FIG. 10  is an expanded lateral sectional view of two adjacent liner assemblies; 
           [0036]      FIG. 11  is an expanded perspective view of an overlapping interface between two adjacent liner assemblies; 
           [0037]      FIG. 12  is an expanded lateral sectional view of two adjacent liner assemblies; 
           [0038]      FIG. 13  is a forward view of two adjacent combustor sections facing a bulkhead heat shield illustrating cooling flow according to one non-limiting embodiment; 
           [0039]      FIG. 14  is a forward view of a combustor section facing a bulkhead heat shield illustrating cooling flow according to another non-limiting embodiment; 
           [0040]      FIG. 15  is a forward view of a combustor section facing a bulkhead heat shield illustrating cooling flow according to another non-limiting embodiment; and 
           [0041]      FIG. 16  is an expanded lateral sectional view of two adjacent liner assemblies. 
       
    
    
     DETAILED DESCRIPTION 
       [0042]      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 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion 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 the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”). 
         [0043]    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 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  (“LPC”) and a low pressure turbine  46  (“LPT”). The inner shaft  40  drives the fan  42  directly or through a geared architecture  48  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. 
         [0044]    The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  (“HPC”) and high pressure turbine  54  (“HPT”). A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  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. 
         [0045]    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 turbines 54, 46 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 bearing structures  38  within the static structure  36 . It should be understood that various bearing structures  38  at various locations may alternatively or additionally be provided. 
         [0046]    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 low pressure compressor  44  and low pressure turbine  46  and render increased pressure in a fewer number of stages. 
         [0047]    A pressure ratio associated with the low pressure turbine  46  is pressure measured prior to the inlet of the low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  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 low pressure compressor  44 , and the low pressure turbine  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. 
         [0048]    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. 
         [0049]    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 (“T”/518.7) 0.5  in which “T” represents the ambient temperature in degrees Rankine. 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). 
         [0050]    With reference to  FIG. 2 , the combustor  56  generally includes an outer combustor liner assembly  60 , an inner combustor liner assembly  62  and a diffuser case module  64 . The outer combustor liner assembly  60  and the inner combustor liner assembly  62  are spaced apart such that a combustion chamber  66  is defined therebetween. The combustion chamber  66  is generally annular in shape. 
         [0051]    The outer combustor liner assembly  60  is spaced radially inward from an outer diffuser case  64 -O 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 -I 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. 
         [0052]    The combustor liner assemblies  60 ,  62  contain the combustion products for direction toward the turbine section  28 . Each combustor liner assembly  60 ,  62  generally includes a respective support shell  68 ,  70  which supports one or more heat shields  72 ,  74  mounted to a hot side of the respective support shell  68 ,  70 . Each of the heat shields  72 ,  74  may be generally rectilinear and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array. In one disclosed non-limiting embodiment, the liner array includes a multiple of forward heat shields  72 A and a multiple of aft heat shields  72 B that are circumferentially staggered to line the hot side of the outer shell  68  (also shown in  FIG. 3 ). A multiple of forward heat shields  74 A and a multiple of aft heat shields  74 B are circumferentially staggered to line the hot side of the inner shell  70  (also shown in  FIG. 3 ). 
         [0053]    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 an annular hood  82 , a bulkhead assembly  84 , a multiple of fuel nozzles  86  (one shown) and a multiple of fuel nozzle pre-swirlers  90  (one shown). Each of the fuel nozzle pre-swirlers  90  is circumferentially aligned with one of the hood ports  94  to project through the bulkhead assembly  84 . Each bulkhead assembly  84  includes a bulkhead support shell  96  secured to the combustor liner assemblies  60 ,  62 , and a multiple of circumferentially distributed bulkhead heat shields  98  secured to the bulkhead support shell  96  around the central opening  92 . 
         [0054]    The annular hood  82  extends radially between, and is secured to, the forwardmost ends of the combustor liner assemblies  60 ,  62 . The annular hood  82  includes a multiple of circumferentially distributed hood ports  94  that accommodate the respective fuel nozzle  86  and introduce air into the forward end of the combustion chamber  66  through a central 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 central opening  92  within the respective fuel nozzle guide  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 chemical 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 the heat shields  72 ,  74  to mount the heat shields  72 ,  74  to the respective support shells  68 ,  70  with fasteners  102  such as nuts (also shown in  FIG. 3 ). That is, the studs  100  project rigidly from the heat shields  72 ,  74  and through the respective support shells  68 ,  70  to receive the fasteners  102  at a threaded distal end section thereof. 
         [0058]    A multiple of cooling impingement holes  104  penetrate through the support shells  68 ,  70  to allow air from the respective annular plenums  76 ,  78  to enter cavities  106 A,  106 B (also shown in  FIG. 3 ) formed in the combustor liner assemblies  60 ,  62  between the respective support shells  68 ,  70  and heat shields  72 ,  74 . The cooling impingement holes  104  are generally normal to the surface of the heat shields  72 ,  74 . The air in the cavities  106 A,  106 B provides backside impingement cooling of the heat shields  72 ,  74  that is generally defined herein as heat removal via internal convection. 
         [0059]    A multiple of cooling film holes  108  penetrate through each of the heat shields  72 ,  74 . The geometry of the film holes, e.g, diameter, shape, density, surface angle, incidence angle, etc., as well as the location of the holes with respect to the high temperature main flow also contributes to effusion film cooling. The combination of impingement holes  104  and film holes  108  may be referred to as an Impingement Film Floatwall liner assembly. 
         [0060]    The cooling film holes  108  allow the air to pass from the cavities  106 A,  106 B defined in part by a cold side  110  of the heat shields  72 ,  74  to a hot side  112  of the heat shields  72 ,  74  and thereby facilitate the formation of a film of cooling air along the hot side  112 . The cooling film holes  108  are generally more numerous than the impingement holes  104  to promote the development of a film cooling along the hot side  112  to sheath the heat shields  72 ,  74 . Film cooling as defined herein is the introduction of a relatively cooler airflow at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the immediate region of the airflow injection as well as downstream thereof. 
         [0061]    A multiple of dilution holes  116  penetrate through both the respective support shells  68 ,  70  and heat shields  72 ,  74  along an axis that could be common or uncommon as indicated in  FIG. 5 . For example only, in a Rich-Quench-Lean (R-Q-L) type combustor, the dilution holes  116  are located downstream of the forward assembly  80  to quench the hot gases by supplying cooling air into the combustor. The hot combustion gases slow towards the dilution holes  116  and may form a stagnation point at the leading edge which becomes a heat source. At the trailing edge of the dilution hole, due to interaction with dilution jet, hot gases form a standing vortex pair that may also become a heat source. 
         [0062]    With reference to  FIG. 6 , a lateral cross-section of the support shells  68 ,  70  and heat shields  72 ,  74  with their respective cavities  106 A,  106 B are illustrated with respect to the combustion chamber  66 . Although only one of the support shells  68 ,  70  and heat shields  72 ,  74  is illustrated and described in detail hereafter, it should be understood that each of the support shells  68 ,  70  and heat shields  72 ,  74  are generally the same and need not be described in detail herein. 
         [0063]    An inner surface  120  of each support shell  68 ,  70  defines a non-planar profile  122  such as a hyperbolic or catenary profile that faces an inner surface  124  of the heat shields  72 ,  74  within the respective cavities  106 A,  106 B. The inner surface  120  of each support shell  68 ,  70  defines a relatively thin cavity zone  126  along a central portion of each combustor section  128  with respect to the inner surface  124  of the heat shields  72 ,  74 . That is, the relatively thin cavity zone  126  is defined generally parallel to the engine axis A and is flanked by relatively thicker cavity zones  128  of each combustor section  130  ( FIG. 5 ). 
         [0064]    With Reference to  FIG. 7 , the convergent support shell  68 ,  70  profile may be the catenary profile is defined by a hyperbolic cosine function, cos h, which provides an approximate 4.5 inlet-to-exit area ratio. The inlet-to-exit area ratio forces a flow acceleration at an end of a circumferential convergent flow section. A corresponding increase in Reynolds number facilitates higher internal heat transfer coefficients for cooling. 
         [0065]    With reference to  FIG. 8 , the relatively thicker cavity zones  128  receive airflow from the impingement holes  104 . The airflow within the cavities  106 A,  106 B is from the relatively thicker cavity zones  128  toward the relatively thin cavity zone  126  to define the circumferential convergent flow section. That is, the airflow is generally in the circumferential direction rather than axial direction. 
         [0066]    In one disclosed non-limiting embodiment, the impingement holes  104  direct airflow onto a multiple of pin fins  132 . The pin fins  132  in one example, may be diamond shaped pins that are approximately ½-¾ the height between the inner surfaces  120 ,  124  in the relatively thicker cavity zones  128 . It should be appreciated that other heights may be provided. 
         [0067]    Inboard of the multiple of pin fins  132 , a multiple of hemispherical nubbins  134  are located toward an exit splitter  136 . In one disclosed non-limiting embodiment, the hemispherical nubbins  134  are of the same diameter but are progressively deeper into the inner surface  124 . That is, centers of the respective spheres which in one disclosed non-limiting embodiment define the hemispherical nubbins  134  are progressively deeper into the combustion chamber  66 . In another disclosed non-limiting embodiment, the hemispherical nubbins  134 - 1  are progressively smaller diameters toward the exit splitter  136  ( FIG. 9 ). The hemispherical nubbins  134 ,  134 - 1  allow for less pressure resistance (less friction) that facilitates convergent flow channel acceleration capabilities. The hemispherical nubbins  134 ,  134 - 1  reduce the frictional drag resistance to the cooling flow yet augment cooling of the inner surface  124 . It should be appreciated that the hemispherical nubbins  134 ,  134 - 1  may be arranged in various patterns. 
         [0068]    The exit splitter  136  is zigzag in shape along the axis A such that a film hole  108  may be located in a valley  138  on each side of the exit splitter  136 . As defined herein “zigzag” includes, but is not limited to, any serpentine, saw tooth or non-straight wall. The exit splitter  136  also forms a base for a frustro-conical stud  100  (only one shown). The stud  100  is received within a corresponding aperture  140  in the heat shields  72 ,  74 , such that as the nut  102  is tightened down on a threaded interface  142 , the aperture  140  seals and tightens onto the frustro-conical stud portion  144  ( FIG. 6 ). 
         [0069]    With reference to  FIG. 10  the threaded interface  144  also forces sets of interleaved cantilevered members  146 ,  148  along each edge  150 ,  152  of the heat shields  72 ,  74  to be forced together to facilitate a seal between each adjacent combustor section  130 - 1 ,  130 - 2  ( FIG. 11 ). It should be appreciated that the frustro-conical studs  100  may alternatively or additionally located in other locations such as along the edges  150 ,  152 . The interleaved cantilevered members  146 ,  148  react the force applied to the frustro-conical stud  100  to minimize leakage. It should be appreciated that the interleaved cantilevered members  146 ,  148  may be of an L-shape, J-shape or other hook-like shape. 
         [0070]    With reference to  FIG. 12 , the film holes  108  along edge  150  of one combustor section  130 - 1  are directed toward edge  152  of the adjacent combustor section  130 - 2  and vice-versa. The cross-flow from the film holes  108  along edges  150 ,  152  protect the edges  150 ,  152  and further facilitate a seal between the interleaved cantilevered members  146 ,  148 . 
         [0071]    The frustro-conical stud  100  and interleaved cantilevered members  146 ,  148  facilitate a relatively higher pressure within the cavities  106 A,  106 B. In one example, an equal number of impingement holes  104  and film holes  108  are located in each combustor section  130  to provide a approximately 50:50 pressure split as compared to a more conventional 80:20 pressure split with approximately half the number of impingement holes  104  compared to the film holes  108 . The 50:50 pressure split permits a relatively higher pressure within the cavities  106 A,  106 B thereby permitting a relatively smaller number of holes and thereby a more efficient usage of air by spacing impingement holes  104  and film holes  108  further apart. Reduced reaction flame temperatures are also avoid local stoichiometric conditions and thereby reduce NOx formation 
         [0072]    With reference to  FIG. 13 , the film holes  108  adjacent to the exit splitter  136  may be directed across an interface  154  between circumferentially distributed bulkhead heat shields  98 . That is, the film holes  108  along one side of the exit splitter  136  are directed toward the opposite side and vice-versa. Such an arrangement may be advantageous when the fuel nozzle pre-swirlers  90  are axially displaced from the film holes  108 . 
         [0073]    With reference to  FIG. 14 , in another disclosed non-limiting embodiment, the film holes  108  on both sides of the zigzag exit splitter  136  through the heat shields  72 ,  74  are directed in a direction in coordination with the rotational direction of the fuel nozzle pre-swirlers  90 . Such an arrangement may be advantageous when the fuel nozzle pre-swirlers  90  are positioned relatively close to the film holes  108 . It should be appreciated that the rotational direction may be clockwise or counter-clockwise. 
         [0074]    With reference to  FIG. 15 , in another disclosed non-limiting embodiment, the film holes  108  on both sides of the zigzag exit splitter  136  through the heat shields  72 ,  74  are directed in a direction in-line the rotational direction of the fuel nozzle pre-swirlers  90 . 
         [0075]    With reference to  FIG. 16 , the non-planar profile  122  may include pre-drilled apertures  156  located in potential hot spots. These apertunes  156  are not initially drilled completely through the support shell  68 ,  70 . That is the pre-drilled apertures  156  are placed in the convergent section close to an area where hot-spots may occur. Should the hot-spot prediction be realized, then apertures  156  are drilled completely through the support shell  68 ,  70  to supply refresher air into the convergent section pre-drilled apertures  156 . This will effectively address the hot-spot by maintaining the coolant heat pick-up low; while introducing more convective flow into the circuit. Furthermore, even if not drilled completely through, the pre-drilled apertures  156  provide weight reduction. 
         [0076]    It should be understood 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. 
         [0077]    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. 
         [0078]    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 in the illustrated embodiment, other arrangements will benefit herefrom. 
         [0079]    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. 
         [0080]    The foregoing description is exemplary rather than defined by the limitations 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 understood 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.