Abstract:
A liner for a combustor of a turbine engine includes a cooling feature which projects from a backside and an effusion hole that communicates through the liner

Description:
BACKGROUND 
       [0001]    The present disclosure relates to a gas turbine engine and, more particularly, to combustor liners with effusion cooling and backside features. 
         [0002]    Gas turbine engines, such as those powering modern commercial and military aircraft, include a compressor for pressurizing 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. 
         [0003]    The desire for combustors that can survive high temperatures yet use less cooling air requires improved cooling efficiency. 
       SUMMARY 
       [0004]    A liner of a combustor for a turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a backside, a cooling feature projecting from the backside, and an effusion hole that communicates through the liner. 
         [0005]    In a further embodiment of the foregoing embodiment, the cooling feature includes a trip strip. 
         [0006]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a pyramid pin fin. 
         [0007]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a three-sided pyramid pin fin. 
         [0008]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a conical pyramid pin fin. 
         [0009]    In a further embodiment of any of the foregoing embodiments, the effusion hole penetrates the cooling feature. 
         [0010]    In a further embodiment of any of the foregoing embodiments, the effusion hole defines an angle less than or equal to ninety (90) degrees with respect to a face of the liner. 
         [0011]    In a further embodiment of any of the foregoing embodiments, the effusion hole is proximate and surrounds an opening through the liner. 
         [0012]    A combustor of a turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a liner with a cooling feature on a backside thereof and a multiple of effusion holes through. 
         [0013]    In a further embodiment of the foregoing embodiment, the cooling feature includes a pin. 
         [0014]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a pyramid pin fin. 
         [0015]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a three-sided pyramid pin fin. 
         [0016]    In a further embodiment of any of the foregoing embodiments, the cooling feature includes a conical pyramid pin fin. 
         [0017]    In a further embodiment of any of the foregoing embodiments, the effusion hole penetrates through at least one of the multiple of cooling features. 
         [0018]    In a further embodiment of any of the foregoing embodiments, the at least one of the multiple of effusion holes defines an angle with respect to a face of the liner. 
         [0019]    In a further embodiment of any of the foregoing embodiments, the effusion hole is defined adjacent to an opening through the liner. 
         [0020]    A combustor of a turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a bulkhead liner with a cooling feature on a backside thereof and a multiple of effusion holes therethrough, the multiple of effusion holes surround an opening through the bulkhead liner. 
         [0021]    In a further embodiment of the foregoing embodiment, the axis defined by a fuel injector passes through the opening. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]    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: 
           [0023]      FIG. 1  is a schematic cross-section of a gas turbine engine; 
           [0024]      FIG. 2  is a perspective partial sectional view of an annular combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0025]      FIG. 3  is an exploded view of a forward assembly of the combustor; 
           [0026]      FIG. 4  is a backside view of a liner according to one disclosed non-limiting embodiment; 
           [0027]      FIG. 5  is an expanded view of a portion of the backside of the combustor panel of  FIG. 4 ; 
           [0028]      FIG. 6  is a backside view of a combustor panel according to another disclosed non-limiting embodiment; 
           [0029]      FIG. 7  is a backside view of a combustor panel according to another disclosed non-limiting embodiment; 
           [0030]      FIG. 8  is a backside view of a combustor panel according to another disclosed non-limiting embodiment; 
           [0031]      FIG. 9  is a backside view of a combustor panel according to another disclosed non-limiting embodiment; 
           [0032]      FIG. 10  is a backside view of a combustor panel according to another disclosed non-limiting embodiment; and 
           [0033]      FIG. 11  is a backside view of a combustor panel according to another disclosed non-limiting embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      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 such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT. 
         [0035]    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. 
         [0036]    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. 
         [0037]    Core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed with the fuel and burned in 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 spool  30  and high spool  32  in response to the expansion. 
         [0038]    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. 
         [0039]    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. 
         [0040]    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 5 (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. 
         [0041]    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. 
         [0042]    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). 
         [0043]    With reference to  FIG. 2 , the combustor  56  generally includes a combustor outer wall  60  and a combustor inner wall  62 . The outer wall  60  and the inner wall  62  are spaced inward from a respective outer and inner wall of a diffuser case module  64  such that a chamber  66  is defined therebetween. The chamber  66  is generally annular in shape and is defined between the walls  60 ,  62 . 
         [0044]    The outer wall  60  and the diffuser case module  64  define an annular outer plenum  76  and the inner wall  62  and the diffuser case module  64  define an annular inner 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. 
         [0045]    Each wall  60 ,  62  generally includes a respective support shell  68 ,  70  that supports one or more respective liners  72 ,  74  mounted to a hot side of the respective support shells  68 ,  70 . The liners  72 ,  74  define a liner array that may be generally annular in shape. Each of the liners  72 ,  74  may be generally rectilinear and manufactured of for example, a nickel based super alloy or ceramic material. 
         [0046]    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 guides  90  (one shown) that define a central opening  92 . The annular hood  82  extends radially between, and may be secured to, the forwardmost ends of the liners  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 . Each fuel nozzle  86  may be secured to the outer case module  64  to project through one of the hood ports  94  and through the central opening  92  within the respective fuel nozzle guide  90 . 
         [0047]    Each of the fuel nozzle guides  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 liners  60 ,  62 , and a multiple of circumferentially distributed liners  98  secured to the bulkhead support shell  96  around the central opening  92  (also Shown in  FIG. 3 ). Each of the fuel nozzle guides  90  is dimensioned to mount a respective the fuel nozzle  86 . 
         [0048]    The forward assembly  80  introduces primary core combustion air into the forward end 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 surrounding structure generate a swirling, intimately blended fuel-air mixture that supports combustion in the forward section of the combustion chamber  66 . 
         [0049]    With reference to  FIG. 4 , each of the liners  72 ,  74 , and/or bulkhead-type liner  98  (shown) include a multiple of support studs  100 , cooling features  102  which project from a backside of the of the liners  72 ,  74 , and/or  98 , and effusion holes  104  (also shown in  FIG. 5 ). In this disclosed non-limiting embodiment, a bulkhead-type liner or heat shield  98  is illustrated, however, the cooling features  102  and the effusion holes  104  are readily applicable to liners  72 ,  74  (FIGS.  2  and  6 - 11 ). It should be further understood that although a significant number of cooling features  102  and effusion holes  104  are illustrated, any number of cooling features  102  and/or effusion holes  104 —including singles—may alternatively be provided. 
         [0050]    The effusion holes  104  generally surround a fuel nozzle opening  106  and in some instances extend directly through the cooling features  102  ( FIG. 5 ). It should be appreciated that the effusion holes  104  may be bored through the cooling features  102  or designed to penetrate areas without the cooling features  102 . 
         [0051]    With reference to  FIG. 6 , the effusion holes  104  define an angle a with respect to a face CF of the liners  72 ,  74 . In the disclosed non-limiting embodiment, the angle a with respect to the face CF may be approximately thirty (30) degrees and oriented along a flow of the combustion gases (illustrated schematically by arrow C). This facilitates the optimization of backside cooling with the benefits of effusion panel cooling as well as impingement cooling. In this disclosed non-limiting embodiment, and those that follow the liners  72 ,  74  are illustrated, however, the cooling features  102  and the effusion holes  104  disclosed are readily applicable to bulkhead liners  98  ( FIGS. 4-5 ). 
         [0052]    With reference to  FIG. 7 , a liner  72 A,  74 A of another disclosed non-limiting embodiment includes a multiple of non-linear trip strips  110 , a multiple of pins  112  and effusion holes  104 . The multiple of pins  112  in the disclosed non-limiting embodiment are square shaped pins, however, any shape may be provided. 
         [0053]    With reference to  FIG. 8 , a liner  72 B,  74 B of another disclosed non-limiting embodiment includes a multiple of substantially linear trip strips  114 , a multiple of pin fins  116  and the effusion holes  104 . 
         [0054]    With reference to  FIG. 9 , a liners  72 C,  74 C of another disclosed non-limiting embodiment includes a multiple of pyramid pin fins  118  and the effusion holes  104 . In this non-limiting embodiment, each of the effusion holes  104  are aligned with a row of the multiple of pyramid pin fins  118  and need not pass therethrough. It should be appreciated that the effusion holes  104  may alternatively pass through one or more of the multiple of pyramid pin fins  118 . 
         [0055]    With reference to  FIG. 10 , liners  72 D,  74 D of another disclosed non-limiting embodiment includes a multiple of three-sided pyramid pin fins  120  and the effusion holes  104 . In this non-limiting embodiment, each of the effusion holes  104  are aligned with a row of the multiple of three-sided pyramid pin fins  120  and need not pass therethrough. 
         [0056]    With reference to  FIG. 11 , liners  72 C,  74 C of another disclosed non-limiting embodiment includes a multiple of conical pin fins  122  and the effusion holes  104 . 
         [0057]    It should be appreciated that various combinations, types and sizes of cooling features or other heat transfer augmenting geometries may be utilized in combination with effusion holes to achieve maximum cooling with a given amount of cooling air. 
         [0058]    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. 
         [0059]    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. 
         [0060]    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. 
         [0061]    Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0062]    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.