Abstract:
A combustor component of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a heat shield panel. The heat shield panel defines a bend and a microcircuit flow path within a thickness of the heat shield panel. The microcircuit flow path includes an inlet and an outlet radially outward of the inlet. The microcircuit flow path at the bend is positioned radially between the inlet and the outlet, and the microcircuit flow path follows the bend. A method of cooling a combustor of a gas turbine engine is also disclosed.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation of U.S. application Ser. No. 13/193,696, filed Jul. 29, 2011. 
     
    
     BACKGROUND 
       [0002]    The present disclosure relates to a combustor, and more particularly to a cooling arrangement therefor. 
         [0003]    Gas turbine combustors have evolved to full hoop shells with attached heat shield combustor liner panels. The liner panels may have relatively low durability due to local hot spots that may cause high stress and cracking. Hot spots are conventionally combated with additional cooling air, however, this may have a potential negative effect on combustor emissions, pattern factor, and profile. 
         [0004]    Current combustor field distresses indicate hot spots at junctions and lips. Hot spots may occur at front bulkhead panels and, in some instances, field distress propagates downstream towards the front liner panels. The distress may be accentuated in local regions where dedicated cooling is restricted due to space limitations. Hot spots may also appear in regions downstream of diffusion quench holes. In general, although effective, a typical combustor chamber environment includes large temperature gradients at different planes distributed axially throughout the combustor chamber. 
       SUMMARY 
       [0005]    A combustor component of a gas turbine engine according to an exemplary aspect of the present disclosure includes a heat shield panel. The heat shield panel defines a bend and a microcircuit flow path within a thickness of the heat shield panel. The microcircuit flow path includes an inlet and an outlet radially outward of the inlet. The microcircuit flow path at the bend is positioned radially between the inlet and the outlet, and the microcircuit flow path follows the bend. 
         [0006]    A method of cooling a combustor of a gas turbine engine according to an exemplary aspect of the present disclosure includes providing a bulkhead heat shield panel defining microcircuit flow path within a thickness of the heat shield panel, and communicating a cooling flow through a bend defined by the microcircuit flow path. The microcircuit flow path includes an inlet and an outlet radially outward of the inlet, the microcircuit flow path at the bend being positioned radially between the inlet and the outlet, and the microcircuit flow path follows the bend. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]    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: 
           [0008]      FIG. 1  is a schematic cross-section of a gas turbine engine; 
           [0009]      FIG. 2  is a perspective partial sectional view of an exemplary annular combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0010]      FIG. 3  is a cross-sectional view of an exemplary combustor that may be used with the gas turbine engine; 
           [0011]      FIG. 4  is a cross-sectional view of the exemplary combustor of  FIG. 3  illustrating a multiple of flows therein; 
           [0012]      FIG. 5  is an expanded cross-sectional view of a bulkhead heat shield panel of the exemplary combustor; 
           [0013]      FIG. 6  is an expanded cross-sectional view of a micro-circuit within the bulkhead heat shield panel; 
           [0014]      FIG. 7  is an expanded transverse cross-sectional view of the micro-circuit of  FIG. 6 ; 
           [0015]      FIG. 8  is an expanded transverse cross-sectional view of the micro-circuit of  FIG. 6  illustrating an in-plane velocity profile; 
           [0016]      FIG. 9  is an expanded cross-sectional view of the micro-circuit of  FIG. 6  illustrating a main velocity profile; 
           [0017]      FIG. 10  is an expanded cross-sectional view of a front heat shield panel of the exemplary combustor; 
           [0018]      FIG. 11  is a planar view of the micro-circuit of  FIG. 6 ; 
           [0019]      FIG. 12  is a sectional view of the micro-circuit of  FIG. 6 ; 
           [0020]      FIG. 13  is an expanded cross-sectional view of a dilution quench section of heat shield panel; 
           [0021]      FIG. 14  is an expanded perspective view of a flow swirler located within a dilution hole; and 
           [0022]      FIG. 15  is an expanded sectional view of the flow swirler. 
       
    
    
     DETAILED DESCRIPTION 
       [0023]      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 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 gas turbine engine 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. 
         [0024]    The 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. 
         [0025]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . 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. 
         [0026]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel within the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines 46, 54 rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0027]    With reference to  FIG. 2 , the combustor  56  generally includes an outer combustor liner  60  and an inner combustor liner  62 . The outer combustor liner  60  and inner combustor liner  62  are spaced inward from a combustor case  64  such that a combustion chamber  66  is defined between combustor liners  60 ,  62 . The combustion chamber  66  may be generally annular in shape. 
         [0028]    The outer combustor liner  60  and the combustor case  64  define an outer annular passageway  76  and the inner combustor liner  62  and the combustor case  64  define an inner annular passageway  78 . It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner panel arrangements will also benefit herefrom. It should be further understood that the cooling flow paths are but an illustrated embodiment and should not be limited only thereto. 
         [0029]    With reference to  FIG. 3 , the combustor liners  60 ,  62  contain and direct the combustion products to the turbine section  28 . Each combustor liner  60 ,  62  generally includes a support shell  68 ,  70  which supports one or more respective heat shield panels  72 ,  74  mounted to a hot side of the respective shell  68 ,  70 . The heat shield panels  72 ,  74  define a heat shield panel array which may be associated with each fuel injector  82 . 
         [0030]    In the disclosed non-limiting embodiment, the heat shield panel array includes forward heat shield panels  72 F and aft heat shield panels  72 A that line the hot side of the outer shell  68  and forward heat shield panels  74 F and aft heat shield panels  74 A that line the hot side of the inner support shell  70 . Fastener assemblies F such as studs and nuts may be used to connect each of the heat shield panels  72 ,  74  to the respective outer and inner support shells  68 ,  70  to provide a floatwall type array. Each of the heat shield panels  72 A,  72 F,  74 A,  74 F may be generally rectilinear and manufactured of, for example, a nickel based super alloy or ceramic material. It should be understood that various numbers, types, and array arrangements of heat shield panels may alternatively or additionally be provided. 
         [0031]    The heat shield panel array may also include a heat shield bulkhead panel  80  that is radially arranged and generally transverse to the heat shield panels  72 ,  74 . Each bulkhead panel  80  surrounds the fuel injector  82  which is mounted within a dome  69  to connect the respective outer and inner support shells  68 ,  70 . Each bulkhead panel  80  may define a bend to provide a transition toward the forward heat shield panels  72 F,  74 F. 
         [0032]    With reference to  FIG. 4 , a cooling arrangement disclosed herein generally includes cooling holes  84 , film cooling holes  86 , dilution holes  88  and refractory metal core (RMC) microcircuits  90  which receive secondary cooling air from passageways  76 ,  78  (shown in  FIG. 3 ). The impingement cooling holes  84  penetrate through the outer and inner support shells  68 ,  70  to communicate the secondary cooling air, into the space between the inner and outer support shells  68 ,  70  and the respective heat shield panels  72 ,  74  to provide backside cooling thereof. The film cooling holes  86  penetrate through each of the heat shield panels  72 ,  74  to promote the formation of a film of cooling air for effusion cooling. Each dilution hole  88  penetrates both the inner and outer support shells  68 ,  70  and the respective heat shield panels  72 ,  74  along a common dilution hole axis D to inject dilution air which facilitates combustion to release additional energy from the fuel. 
         [0033]    The RMC microcircuits  90  may be selectively formed within the heat shield panels  72 ,  74  and the bulkhead panels  80  through a refractory metal core process. Refractory metal cores (RMCs) are typically metal based casting cores usually composed of molybdenum with a protective coating. The refractory metal provides more ductility than conventional ceramic core materials while the coating—usually ceramic—protects the refractory metal from oxidation during a shell fire step of the investment casting process and prevents dissolution of the core from molten metal. 
         [0034]    With reference to  FIG. 5 , the cooling arrangement disclosed in the illustrated non-limiting embodiment locates RMC microcircuits  90 A within the bulkhead panels  80 . Split lines from RMC tabs formed in the RMC process may be used to define one or more inlets  92  and/or exits  94  ( FIG. 6 ). Although the RMC microcircuit  90 A is illustrated as generally circular in cross-section ( FIG. 7 ), it should be understood that various profile cross-sections may alternatively or additionally be provided. 
         [0035]    A bend  98  is defined within each RMC microcircuit  90 A within the bulkhead panel  80  such that the Dean&#39;s effect becomes a driving force to establish in-plane re-circulating flows ( FIG. 8 ) which facilitate enhanced internal cooling of the bulkhead panels  80  as the cooling flow is essentially biased toward the hot side thereof ( FIG. 9 ). 
         [0036]    With reference to  FIG. 10 , the cooling arrangement may continue through RMC microcircuits  90 B located within the forward panels  72 F,  74 F. Each RMC microcircuit  90 B within the forward panels  72 F,  74 F include a multiple of internal features  100  positioned in a flow path of the microcircuits  90 B ( FIG. 11 ) and which may be repeated tangentially and axially throughout the heat shield panels  72 ,  74 . 
         [0037]    With reference to  FIG. 11 , the internal features  100  are arranged as a pair of islands positioned in a flow path of the microcircuit  90 B. The islands or internal features  100  define metering sections  102  which operate as flow restrictions to control secondary cooling air flow and optimize cooling efficiency between an inlet  104  and an exit  106 . The metering sections  102  facilitate optimum convective efficiency as a measure of heat pick-up, as the higher heat pick-up leads to higher cooling efficiency. It should be understood that various shapes of internal features shaped other than the illustrated islands may alternatively or additionally be provided. 
         [0038]    The inlets  104  and exits  106  may provide for film cooling, impingement cooling or combinations thereof. The exits  106  may alternate between film and impingement as required for management of the external thermal load. In one disclosed non-limiting embodiment, exits  106  are in communication with the microcircuit  90 B and are arranged adjacent to film cooling holes  86  which penetrate through each of the heat shield panels  72 ,  74  to promote the formation of a film of cooling air for effusion cooling, and in another example are arranged adjacent to impingement cooling holes  84  (shown in  FIG. 12 ). In addition, a wall depression  108  is located on a hot side of the heat shield panels  72 ,  74  downstream of each exit  106 ,  86  ( FIG. 12 ) to facilitate film wall attachment through the Coanda effect. That is, the microcircuit  90 B provides for effective convective efficiency supplemented alternatively by either impingement or film cooling with wall depressions  108  for film cooling protection. 
         [0039]    With reference to  FIG. 13 , a flow swirler  110  is located within each dilution hole  88  to combine dilution flow with RMC microcircuit  90 B to vector the dilution flow. In the disclosed non-limiting embodiment, the flow swirler  110  is located upstream of the combustor lean zone. The flow swirler  110  defines an axial flow path d for the dilution flow and a transverse flow path for receipt of the RMC microcircuit flow t ( FIG. 14 ). Vanes  112  within the flow swirler  110  combine the flows d, t to vector the dilution flow d tangentially and accelerate the RMC microcircuit flow t ( FIG. 15 ). The sink pressures for the RMC microcircuits are also lowered as a consequence of the high velocity flow swirler  110  to reduce cooling flux by approximately 40% over conventional levels. 
         [0040]    The RMC microcircuits  90  provide effective cooling to address gas temperature variations inside the combustor chamber; enhance cooling through flow distribution with heat transfer enhancement features while maintaining increased film coverage and effectiveness throughout the combustor chamber; improve combustor durability by optimum distribution of cooling circuits; and facilitate lower emissions and improved turbine durability. 
         [0041]    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. 
         [0042]    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. 
         [0043]    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. 
         [0044]    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.