Patent Publication Number: US-8978385-B2

Title: Distributed cooling for gas turbine engine combustor

Description:
BACKGROUND 
     The present disclosure relates to a combustor, and more particularly to a cooling arrangement therefor. 
     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. 
     Current combustor field distresses indicate hot spots at junctions and lips. Hot spots may occur at front heat shield 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 
     A combustor component of a gas turbine engine according to an exemplary aspect of the present disclosure includes a liner panel with a refractory metal core (RMC) microcircuit. 
     A method of cooling a combustor of a gas turbine engine according to an exemplary aspect of the present disclosure includes self regulating a cooling flow through a refractory metal core (RMC) microcircuit within a heat shield. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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: 
         FIG. 1  is a schematic cross-section of a gas turbine engine; 
         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 ; 
         FIG. 3  is a cross-sectional view of an exemplary combustor that may be used with the gas turbine engine; 
         FIG. 4  is an expanded plan view of a microcircuit; 
         FIG. 5  is an expanded cross-sectional view of the microcircuit of  FIG. 5 ; 
         FIG. 6A  is a plan view of a first flow condition within the liner panel; 
         FIG. 6B  is a plan view of a second flow condition within the liner panel; 
         FIG. 7A  is a first example flow distribution which is unbalanced. 
         FIG. 7B  is a second example flow distribution which is unbalanced and the reverse of  FIG. 7A ; 
         FIG. 8  is a flow chart of microcircuit operation; 
         FIG. 9  is a planar view of another microcircuit; and 
         FIG. 10  is a sectional view of the microcircuit of  FIG. 9 . 
     
    
    
     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 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. 
     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. 
     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. 
     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  54 ,  46  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     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 the inner combustor liner  62  are spaced inward from a combustor case  64  such that a combustion chamber  66  is defined there between. The combustion chamber  66  is generally annular in shape and is defined between combustor liners  60 ,  62 . 
     The outer combustor liner  60  and the combustor case  64  define an outer annular passageway  76 . 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 disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto. 
     With reference to  FIG. 3 , the combustor liners  60 ,  62  contain the flame for direction toward the turbine section  28 . Each combustor liner  60 ,  62  generally includes a support shell  68 ,  70  which supports one or more liner panels  72 ,  74  mounted to a hot side of the respective support shell  68 ,  70 . The liner panels  72 ,  74  define a liner panel array which may be generally annular in shape. Each of the liner panels  72 ,  74  may be generally rectilinear and manufactured of, for example, a nickel based super alloy or ceramic material. 
     In the disclosed non-limiting embodiment, the combustor  56  includes a plurality of liner panels  72 ,  74  arranged about a combustor axis C to define an array. A plurality of forward liner panels  72 F and aft liner panels  72 A line the hot side of the outer shell  68 , and forward liner panels  74 F and aft liner panels  74 A line the hot side of the inner shell  70 . Fastener assemblies F such as studs and nuts may be used to connect each of the liner panels  72 ,  74  to the respective inner and outer shells  68 ,  70  to provide a floatwall type array. It should be understood that various numbers, types, and array arrangements of liner panels may alternatively or additionally be provided. 
     The combustor  56  may also include heat shield panels  80  that are radially arranged and generally transverse to the liner panels  72 ,  74 . Each heat shield panel  80  surrounds a fuel injector  82  which is mounted within a dome  69  which connects the respective inner and outer support shells  68 ,  70 . 
     A cooling arrangement disclosed herein may generally include a multiple of impingement cooling holes  84 , film cooling holes  86 , dilution holes  88  and refractory metal core (RMC) microcircuits  90  (illustrated schematically). The impingement cooling holes  84  penetrate through the inner and outer support shells  68 ,  70  to communicate coolant, such as a secondary cooling air, into the space between the inner and outer support shells  68 ,  70  and the respective liner panels  72 ,  74  to provide backside cooling thereof. The film cooling holes  86  penetrate each of the liner panels  72 ,  74  to promote the formation of a film of cooling air for effusion cooling. The dilution holes  88  penetrate both the inner and outer support shells  68 ,  70  and the respective liner panels  72 ,  74  along a common dilution hole axis d to inject dilution air which facilitates combustion and release additional energy from the fuel. 
     Referring to  FIGS. 3-5 , the RMC microcircuits  90  may be selectively formed within the liner panels  72 ,  74  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. The refractory metal core process allows small features to be cast inside internal passages. This, in turn, allows advanced cooling concepts, through the design space with relatively lower cooling flows as compared to current technology cooling flow levels. 
     RMC technology facilitates the manufacture of very small cast features such that the cooling supply flow may be minimized. As the cooling supply flow decreases, it may be beneficial to minimize any flow arrangement that may not operate at the highest level of optimization. Therefore, the design of the RMC microcircuit may beneficially optimize flow distribution by sensing external operating conditions. 
     With reference to  FIG. 4 , an RMC microcircuit  90 A according to one non-limiting embodiment is formed within the liner panel  72 ,  74 . In the disclosed non-limiting embodiment, the height ( FIG. 5 ) of the RMC microcircuit  90 A may be in the range of 0.012-0.025 inches (0.030-0.064 cm) for each location within each liner panel  72 ,  74 . That is, the liner panel  72 ,  74  includes the disclosed internal features which are formed via RMC technology. It should be understood that various heights may alternatively or additionally be provided. 
     Referring to  FIGS. 4 and 5 , the RMC microcircuit  90 A includes a multiple of internal features located within the generally rectilinear liner panel  72 ,  74 . The internal features extend radially between liner sections  75 . The internal features may generally include a semi-circular inlet  92 , a first divergent island  94 A, a second divergent island  94 B, a flow separator island  98 , a first feedback feature  100 A, a second feedback feature  100 B, a first slot exit  102 A and a second slot exit  102 B (also shown in  FIG. 5 ). The feedback features  100 A,  100 B extend from walls  77  that bound the secondary flow. In some examples, the exit slots  102 A,  102 B can be arranged coaxially with an adjacent liner panel  72 ,  74  (shown in  FIG. 5 ). The internal features include an inlet wall  93  having a semi-circular geometry extending from a first wall  95  of the liner panel  72 ,  74  to provide the inlet  92 . An access port  79  (shown in  FIG. 5 ) extends from the liner panel  75  to communicate flow between the inner and outer annular passageways  76 .  78  and the inlet  92 , As shown, the access port  79  extends through the support shell  68 ,  70 . The inlet wall  93  bounds the inlet  92  to direct flow between the inner and outer annular passageways  76 ,  78  and a main flow path or cooling channel  104 . Generally, the first divergent island  94 A, the second divergent island  94 B, the flow separator island  98 , the first feedback feature  100 A, and the—second feedback feature  100 B are structures formed by the RMC microcircuit  90 A which guide and direct the secondary flow as described herein within the cooling channel  104  formed within the liner panel  72 ,  74 . That is, the structures form flows such as a self-regulating feedback which is further describe herein below. The inlet  92 , the first slot exit  102 A and the second slot exit  102 B provide communication into or out of the RMC microcircuit  90 A. That is, the liner panel  72 ,  74 , the inlet  92 , the first slot exit  102 A and the second slot exit  102 B provide communication from within the liner panel  72 ,  74  to the combustor chamber  66 . 
     In this non-limiting embodiment, the semi-circular inlet  92  and the flow separator island  98  are located along an axis P. In some examples, the inlet wall  93  is at least partially coaxial with the divergent islands  94 A.  94 B along the axis P. As shown in  FIG. 4 , the first and second divergent islands  94 A,  94 B extend a distance  97  along the axis P. and the inlet wall  93  and flow separator island are spaced apart a distance  99  along the axis P such that distance  97  is greater than distance  99 . Also as shown in  FIG. 4 . the first and second divergent islands  94 A,  94 B are spaced a distance  101  from the first wall  95 . which is less than a length  103  of the inlet wall  91  In this arrangement, an inlet port  105  defined by the inlet wall  93  extends downstream of a feedback outlet  107  provided by one the first and second divergent islands  94 A,  94 B with respect to the axis P. The first divergent island  94 A may define a location for a dilution hole  88  which extends therethrough. The second divergent island  94 B may define a mount for the fastener F which supports the liner panel  72 ,  74  ( FIG. 5 ). It should be understood that other arrangements of internal features, fastener and hole locations may alternatively or additionally be provided. 
     With reference to  FIG. 6A , a feedback feature  100 A,  100 B may be transverse and extend toward the axis P to facilitate generation of self-regulating feedback loops or flow paths S 1 , S 2 . The semi-circular inlet  92  forces the secondary cooling air S to spread into a cooling channel  104 . The divergent islands  94 A,  94 B are configured to further spread the flow in the channel  104 . As the cooling flow approaches slot exits  102 A,  102 B, the self-regulating feedback flow paths S 1 , S 2  form loops around the respective divergent islands  94 A,  94 B. The first and second feedback loops S 1 , S 2  each include a feedback passage  114  extending between the feedback outlet  107  and a feedback inlet  113  positioned downstream of the feedback outlet  107  (shown in  FIGS. 4 and 6A ). As shown, the feedback outlets and inlets  107 .  113  are defined between the walls  77  and the divergent islands  94 A,  94 B. As shown in  FIG. 4 , each of the feedback features  100 A,  100 B extends radially inward a first distance  115  greater than a second distance  117  defined by each of the feedback inlets  113  to communicate flow from the channel  104  to each of the feedback inlets  113 . The internal features adjust the internal cooling flow characteristics in response to an operating condition as represented graphically by flow distributions at stations (i) and (i+1). 
     If the secondary cooling air S flow velocity is uniform within the channel  104  formed by islands  94 A,  94 B, the self-regulating feedback flows S 1 , S 2  are equivalent, and there is no preferred tendency for the flow of secondary cooling air S to move to either of the exit slots  102 A,  102 B. However, if the secondary cooling air S flow velocity is not uniform, an unbalance between the self-regulating feedback flows S 1 , S 2  will be established to modulate the flow to the respective slot exits  102 A,  102 B ( FIGS. 6A ,  6 B). In  FIG. 6A , an example flow distribution ( FIG. 7A ) is illustrated when the secondary cooling air S flow velocities increase towards the slot exit  102 A (station (i+1)). The reverse occurs in  FIG. 6B  as the main secondary cooling air S flow velocities increases towards the slot exit  102 B (station (i)). This effect attenuates potential hot streaks in the main secondary cooling air S flow through increased film cooling where required ( FIG. 7B ). That is, the self regulating feedback flows S 1 , S 2  sense the effects of the sink pressure changes and influences flow of the main secondary cooling air S distribution to address the fluctuations and balance in a self-regulating manner ( FIG. 8 ). The transfer of flow control is derived from sensing the sink pressure variations at the microcircuit exit. The flow rate within the microcircuit is inversely proportional to the sink pressure variations. As a result, the feedback flow returns to the beginning of the circuit, which then directs the main flow to the flow branch whose exit has a relative higher sink pressure. This provides a self-regulating action in the circuit without any moving parts. 
     With reference to  FIG. 9 , an RMC microcircuit  90 B according to another non- limiting embodiment, formed within the liner panel  72 A,  74 A supplements the internal features as discussed above. The microcircuit  90 B includes a first region  108  and a second region  109  separated by a flow separator island  98 ′. An axis P extends between a first wall  95  and a second wall  111  of the liner panel  72 ,  74 . Cooling enhancement features such as pedestals  106 A, followed by flow straighteners  106 B, are formed in the second region  109  and upstream of slot film cooling openings  110  (also shown in  FIG. 9 ). As shown, the slot film cooling openings  110  include a first opening  110 A located along the axis P and one or more second openings  110 B offset from the axis P. These relatively small cooling enhancement features are structures formed within the second region  109  to further effect the flow and are readily manufactured through refractory metal core technology in a manner commensurate with the islands  94 A,  94 B. Additionally, a multiple of laser holes  112  (illustrated schematically) may be located at strategic locations ahead of relatively larger internal features. 
     In this non-limiting embodiment, the feedback features  100 A′,  100 B′ define a metering area between the internal features  94 A,  94 B and the cooling enhancement features  106 A,  106 B. The indented feedback features  100 A′,  100 B′ also provide a location for a dilution hole  88 ′. The flow separator island  98 ′ may define a mount for the fastener F which supports the liner panel  72 A,  74 A ( FIG. 10 ). 
     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. 
     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. 
     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. 
     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. 
     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.