Patent Publication Number: US-2019186739-A1

Title: Apparatus and method for mitigating particulate accumulation on a component of a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims the benefit of U.S. Provisional Application No. 62/607,606 filed Dec. 19, 2017, which is incorporated herein by reference in its entirety. 
    
    
     BACKGROUND 
     The subject matter disclosed herein generally relates to gas turbine engines and, more particularly, to method and apparatus for mitigating particulate accumulation on cooling surfaces of components of gas turbine engines. 
     In one example, a combustor of a gas turbine engine may be configured and required to burn fuel in a minimum volume. Such configurations may place substantial heat load on the structure of the combustor (e.g., panels, shell, etc.). Such heat loads may dictate that special consideration is given to structures, which may be configured as heat shields or panels, and to the cooling of such structures to protect these structures. Excess temperatures at these structures may lead to oxidation, cracking, and high thermal stresses of the heat shields or panels. Particulates in the air used to cool these structures may inhibit cooling of the heat shield and reduce durability. Particulates, in particular atmospheric particulates, include solid or liquid matter suspended in the atmosphere such as dust, ice, ash, sand and dirt. 
     SUMMARY 
     According to one embodiment, a gas turbine engine component assembly is provided. The gas turbine engine component assembly comprising: a first component having a first surface, a second surface opposite the first surface, a first cooling hole located in a first section of the first component extending from the second surface to first surface, and a second cooling hole located in a second section of the first component extending from the second surface to first surface; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween in fluid communication with the cooling hole for cooling the second surface of the second component; wherein the first cooling hole is configured to direct at least one of the airflow and the particulate to impinge upon the second surface of the second component at first directional flow angle, and wherein the second cooling hole is configured to direct at least one of the airflow and the particulate to impinge upon the second surface of the second component at a second directional flow angle different from the first directional flow angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first cooling hole is configured to direct at least one of the airflow and the particulate to impinge upon the second surface of the second component at a first impingement angle, and wherein the second cooling hole is configured to direct at least one of the airflow and the particulate to impinge upon the second surface of the second component at a second impingement angle different from the first impingement angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that at least one of the first impingement angle and the second impingement angle is non-perpendicular. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that that first cooling hole is formed in the first component with a non-perpendicular primary aperture angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second cooling hole is formed in the first component with a non-perpendicular primary aperture angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first directional flow angle is equivalent to a directional angle of a local cross-flow path within the cooling channel. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second directional flow angle is equivalent to a directional angle of a local cross-flow path within the cooling channel. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second surface of the second component is non-planar to the first surface of the first component. 
     According to another embodiment, a shell of a combustor for use in a gas turbine engine is provided. The shell comprising: a combustion chamber of the combustor, the combustion chamber having a combustion area; a combustion liner having an inner surface, an outer surface opposite the inner surface, a first primary aperture located in a first section of the combustion liner extending from the outer surface to the inner surface through the combustion liner, and a second primary apertures located in a second section of the combustion liner extending from the outer surface to the inner surface through the combustion liner; a heat shield panel interposed between the inner surface of the combustion liner and the combustion area, the heat shield panel having a first surface and a second surface opposite the first surface, wherein the second surface is oriented towards the inner surface, and wherein the heat shield panel is separated from the combustion liner by an impingement cavity, wherein the first primary aperture is configured to direct at least one of the airflow and the particulate to impinge upon the second surface at first directional flow angle, and wherein the second primary aperture is configured to direct at least one of the airflow and the particulate to impinge upon the second surface at a second directional flow angle different from the first directional flow angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first primary aperture is configured to direct at least one of the airflow and the particulate to impinge upon the second surface at a first impingement angle, and wherein the second primary aperture is configured to direct at least one of the airflow and the particulate to impinge upon the second surface at a second impingement angle different from the first impingement angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that at least one of the first impingement angle and the second impingement angle is non-perpendicular. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first primary aperture is formed in the combustion liner with a non-perpendicular primary aperture angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second primary aperture is formed in the combustion liner with a non-perpendicular primary aperture angle. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the first directional flow angle is equivalent to a directional angle of a local cross-flow path within the impingement cavity. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second directional flow angle is equivalent to a directional angle of a local cross-flow path within the impingement cavity. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second surface is non-planar to the inner surface. 
     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, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: 
         FIG. 1  is a partial cross-sectional illustration of a gas turbine engine, in accordance with an embodiment of the disclosure; 
         FIG. 2  is a cross-sectional illustration of a combustor, in accordance with an embodiment of the disclosure; 
         FIG. 3 a    is an enlarged cross-sectional illustration of a heat shield panel and combustion liner of a combustor, in accordance with an embodiment of the disclosure; 
         FIG. 3 b    is a cross-sectional illustration of a particulate collection mitigation system for a combustor of a gas turbine engine, in accordance with an embodiment of the disclosure; 
         FIG. 3 c    is an illustration of a bulkhead portion of a combustion liner for a combustor of a gas turbine engine, in accordance with an embodiment of the disclosure; and 
         FIG. 3 d    is an illustration of a bulkhead portion of a combustion liner for a combustor of a gas turbine engine, in accordance with an embodiment of the disclosure. 
     
    
    
     The detailed description explains embodiments of the present disclosure, together with advantages and features, by way of example with reference to the drawings. 
     DETAILED DESCRIPTION 
     A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures. 
     Combustors of gas turbine engines, as well as other components, experience elevated heat levels during operation. Impingement and convective cooling of panels of the combustor wall may be used to help cool the combustor. Convective cooling may be achieved by air that is channeled between the panels and a liner of the combustor. Impingement cooling may be a process of directing relatively cool air from a location exterior to the combustor toward a back or underside of the panels. 
     Thus, combustion liners and heat shield panels are utilized to face the hot products of combustion within a combustion chamber and protect the overall combustor shell. The space between the combustion liner and the heat shield panel is often called the impingement cavity. The combustion liners may be supplied with cooling air including dilution passages which deliver a high volume of cooling air into a hot flow path. The cooling air may be air from the compressor of the gas turbine engine. The cooling air may impinge upon a back side of a heat shield panel in the impingement cavity that faces a combustion liner inside the combustor. The cooling air may contain particulates, which may collect on the heat shield panels overtime, thus reducing the cooling ability of the cooling air. The collection of particulate on the heat shield panel may be due to aerodynamics within the impingement cavity. Aerodynamics in impingement cavity can be turbulent due to the expansion and mixing of the multitude of impingement airflows. This turbulence leads to locally low velocities, which may contribute to increased rate of dirt deposition on the backside of panels. Embodiments disclosed herein seek to address particulate adherence to the heat shield panels in order to maintain the cooling ability of the cooling air. 
       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 flow path B in a bypass duct, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
     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 speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as 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  300  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . An engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The engine static structure  36  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  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 in the combustor  300 , 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. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio 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. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. 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. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about  0 .8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second ( 350 . 5  m/sec). 
     Referring now to  FIG. 2  and with continued reference to  FIG. 1 , the combustor section  26  of the gas turbine engine  20  is shown. As illustrated in  FIG. 2 , a combustor  300  defines a combustion chamber  302 . The combustion chamber  302  includes a combustion area  370  within the combustion chamber  302 . The combustor  300  includes an inlet  306  and an outlet  308  through which air may pass. The air may be supplied to the combustor  300  by a pre-diffuser  110 . Air may also enter the combustion area  370  of the combustion chamber  302  through other holes in the combustor  300  including but not limited to quench holes  310 , as seen in  FIG. 2 . 
     As shown in  FIG. 2 , compressor air is supplied from a compressor section  24  into a pre-diffuser strut  112 . As will be appreciated by those of skill in the art, the pre-diffuser strut  112  is configured to direct the airflow into the pre-diffuser  110 , which then directs the airflow toward the combustor  300 . The combustor  300  and the pre-diffuser  110  are separated by a shroud chamber  113  that contains the combustor  300  and includes an inner diameter branch  114  and an outer diameter branch  116 . As air enters the shroud chamber  113 , a portion of the air may flow into the combustor inlet  306 , a portion may flow into the inner diameter branch  114 , and a portion may flow into the outer diameter branch  116 . 
     The air from the inner diameter branch  114  and the outer diameter branch  116  may then enter the combustion area  370  of the combustion chamber  302  by means of one or more primary apertures  307  in the combustion liner  600  and one or more secondary apertures  309  in the heat shield panels  400 . The primary apertures  307  and secondary apertures  309  may include nozzles, holes, etc. The air may then exit the combustion chamber  302  through the combustor outlet  308 . At the same time, fuel may be supplied into the combustion chamber  302  from a fuel injector  320  and a pilot nozzle  322 , which may be ignited within the combustion area  370  of the combustion chamber  302 . The combustor  300  of the engine combustion section  26  may be housed within a shroud case  124  which may define the shroud chamber  113 . 
     The combustor  300 , as shown in  FIG. 2 , includes multiple heat shield panels  400  that are attached to the combustion liner  600  (See  FIG. 3 a   ). The heat shield panels  400  may be arranged parallel to the combustion liner  600 . The combustion liner  600  can define circular or annular structures with the heat shield panels  400  being mounted on a radially inward liner and a radially outward liner, as will be appreciated by those of skill in the art. The heat shield panels  400  can be removably mounted to the combustion liner  600  by one or more attachment mechanisms  332 . In some embodiments, the attachment mechanism  332  may be integrally formed with a respective heat shield panel  400 , although other configurations are possible. In some embodiments, the attachment mechanism  332  may be a bolt or other structure that may extend from the respective heat shield panel  400  through the interior surface to a receiving portion or aperture of the combustion liner  600  such that the heat shield panel  400  may be attached to the combustion liner  600  and held in place. The heat shield panels  400  partial enclose a combustion area  370  within the combustion chamber  302  of the combustor  300 . 
     Referring now to  FIGS. 3 a , 3 b , 3 c , and 3 d    with continued reference to  FIGS. 1 and 2 .  FIG. 3 a    illustrates a heat shield panel  400  and a combustion liner  600  of a combustor  300  (see  FIG. 1 ) of a gas turbine engine  20  (see  FIG. 1 ). The heat shield panel  400  and the combustion liner  600  are in a facing spaced relationship.  FIG. 3 b    shows a particulate collection mitigation system  100  for a combustor  300  (see  FIG. 1 ) of a gas turbine engine  20  (see  FIG. 1 ), in accordance with an embodiment of the present disclosure. The heat shield panel  400  includes a first surface  410  oriented towards the combustion area  370  of the combustion chamber  302  and a second surface  420  first surface opposite the first surface  410  oriented towards the combustion liner  600 . The combustion liner  600  has an inner surface  610  and an outer surface  620  opposite the inner surface  610 . The inner surface  610  is oriented toward the heat shield panel  400 . The outer surface  620  is oriented outward from the combustor  300  proximate the inner diameter branch  114  and the outer diameter branch  116 . 
     The combustion liner  600  includes a plurality of primary apertures  307  configured to allow airflow  590  from the inner diameter branch  114  and the outer diameter branch  116  to enter an impingement cavity  390  in between the combustion liner  600  and the heat shield panel  400 . Each of the primary apertures  307  extend from the outer surface  620  to the inner surface  610  through the combustion liner  600 . 
     Each of the primary apertures  307  fluidly connects the impingement cavity  390  to at least one of the inner diameter branch  114  and the outer diameter branch  116 . The primary apertures  307  are configured to direct airflow  590  towards the second surface  420  of the heat shield panel  400  and the directed airflow  590  provides cooling to the heat shield panel  400  when the airflow impinges on the second surface at an impingement point  594 . The airflow  590  may strike or impinge upon the second surface  420  at an impingement angle α 1 , that is conventionally about 90° or about perpendicular. An impingement angle α 1  about equal to 90° may lead to some turbulence of airflow  590  within the impingement cavity  390 , which may lead to collection of particulate  592  on the second surface  420  of the heat shield panel  400 , as described further below. The impingement angle α 1  may be adjusted by the primary aperture angle β 1  of each primary aperture  307  along with the angular orientation of the combustor liner  600  relative to the heat shield panel  400 . 
     The heat shield panel  400  may include one or more secondary apertures  309  configured to allow airflow  590  from the impingement cavity  390  to the combustion area  370  of the combustion chamber  302 . Each of the secondary apertures  309  extend from the second surface  420  to the first surface  410  through the heat shield panel  400 . Airflow  590  flowing into the impingement cavity  390  impinges on the second surface  420  of the heat shield panel  400  at an impingement point  594  and absorbs heat from the heat shield panel  400  as it impinges on the second surface  420 . As seen in  FIG. 3 a   , particulates  592  may accompany the airflow  590  flowing into the impingement cavity  390 . Particulate  592  may include but are not limited to dirt, smoke, soot, volcanic ash, or similar airborne particulate known to one of skill in the art. As the airflow  590  and particulates  592  impinge upon the second surface  420  of the heat shield panel  400 , the pollutant particulate  592  may begin to collect on the second surface  420 , as seen in  FIG. 3 a   . Particulate  592  collecting upon the second surface  420  of the heat shield panel  400  reduces the cooling efficiency of airflow  590  impinging upon the second surface  420 , and thus may increase local temperatures of the heat shield panel  400  and the combustion liner  600 . Particulate  592  collecting upon the second surface  420  of the heat shield panel  400  may potentially create a blockage  593  to the secondary apertures  309  in the heat shield panels  400 , thus reducing airflow  590  into the combustion area  370  of the combustion chamber  302 . The blockage  593  may be a partial blockage or a full blockage. 
     Particulate  592  tends to collect at various collection points along second surface  420  of the heat shield panel  400 . The collection points may include impingement points  594  and impingement flow convergence point  595 . Impingement points  594  are points on the second surface  420  of the heat shield panel  400  where the airflow  590  and particulate  592  from a first primary aperture  307  is directed to impinge upon the second surface of the heat shield panel. Thus, each impingement points  594  is located opposite a primary aperture  307 . When the airflow  590  and particulate  592  hit the second surface  594 , the airflow and particulate  592  are forced to change direction abruptly, thus resulting in a loss of speed. The direction change will be either in a first direction  90  or a second direction  92 . This direction change and loss of speed will result in some particulate  592  being separated from the airflow  590  and the particulates  590  that are separated will collect at the impingement point  594 , as seen in  FIG. 3 a   . The particulate  592  that does not collect at the impingement point  594  will be directed along with the airflow  592  either in a first direction  90  or a second direction  92  until the particulate  592  and airflow  590  converges at a impingement flow convergence point  595  with the particulate  592  and airflow  590  from a second primary aperture  307  adjacent to the first primary aperture  307 , as seen in  FIG. 3 a   . Each impingement flow convergence point  595  may be located about equally between two or more impingement points  594 , as seen in  FIG. 3 a   . At an impingement flow convergence point  595 , the converging particulate  592  and airflow  590  is forced to change direction abruptly for a second time, thus resulting in a loss of speed. The second direction change will be towards the combustion liner  600 . This second direction change and loss of speed will result in some particulate  592  being separated from the airflow  590  and the particulates  590  separated will collect at the impingement flow convergence point  595 , as seen in  FIG. 3   a.    
     The combustion liner  600  may include one or more primary apertures  307  configured to direct at least one of airflow and particulate  592  to a second surface  420  to impinge upon the second surface  420  at an impingement angle α 1  that is non-perpendicular (i.e. the impingement angle is not equal to 90°), as seen in  FIG. 3 b   . In order to produce an impingement angle α 1  that is non-perpendicular, the primary apertures  30  may be formed in the combustor liner  600  with a non-perpendicular primary aperture angle β 1 . The primary aperture angle β 1  may be measured with respect to the inner surface  610 , as seen in  FIG. 3 b   . In an alternative embodiment, in order to produce an impingement angle α 1  that is non-perpendicular, a plane angle γ 1  measured between the inner surface  610  and the second surface  420  may be not equal to 180° (i.e. the second surface  420  is non-planar to the inner surface  610 ). In another alternative embodiment, a supplemental flow directing mechanism may be inserted into the primary aperture  307  to passively and/or actively direct the airflow  590  and/or particles  592  expelled from the primary aperture  307 , thus adjusting the impingement angle α 1 . In an embodiment, the impingement angle α 1  may be oriented such that at least one of the airflow  590  and particulates  592  are directed in a direction of a local cross-flow path D within the impingement cavity  390 , as seen in  FIG. 3 b   . Advantageously by impinging airflow  590  onto the second surface  420  at an angle relative to the second surface  420  that is non-perpendicular the cooling airflow  590  may be directed towards a preferential direction which can minimize the local low velocity regions. 
     A bulkhead portion  700  of the combustion liner  600  may be seen in  FIGS. 3 c  and 3 d   . The bulkhead portion  700  may be located on the forward end of the combustor  300  and includes a through hole  710  configured to fit the combustor inlet  306  and pilot nozzle  322  of the fuel injectors  322 . The combustor panel  600  may be sub-divided into separate sections and each section may include primary apertures  307  configured to direct the airflow  590  and particulate  592  (not shown in  FIG. 3 c   ) at different impingement angles α 1  from each other section. In the example illustrated in  FIG. 3 c   , the combustor panel  600  is sub-divided into 5 separate sections, each having primary apertures  307  configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) at different impingement angles α 1  and/or different directional flow angle Θ 1 . The directional flow angle Θ 1  is the angle that the airflow  590  will be directed across the heat shield panel  400 . The directional flow angle Θ 1  may be measured relative to an axis X 1 . The directional flow angle Θ 1  may be about equal to a local cross-flow path in the impingement cavity  390 . Advantageously, if the directional flow angle Θ 1  the local cross-flow path in the impingement cavity  390 , the impediment of airflow  590  from the primary aperture  307  upon the cross-flow airflow  590  within the impingement cavity will be reduced. 
     In one example, each section may have primary apertures  307  with differing directional flow angles Θ 1  between the sections. In another example, the primary apertures  307  within a section may have differing directional flow angles Θ 1 . In another example, each section may have primary apertures  307  with differing primary aperture angles β 1  between the sections to produce differing impingement angles α 1 . The five sections include a radially outward section  614 , a readily inward section  616 , a first section  618 , a second section  622 , and a center section  624 . 
     In the radially outward section  614 , the primary apertures  307  are configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards a radially outward side  604  of the bulkhead portion  700  of the combustion liner  600 . In an embodiment, the primary apertures  307  in the radially outward section  614  may include a primary aperture angle β 1  configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards the radially outward side  604  of the bulkhead portion  700  of the combustion liner  600   
     In the radially inward section  616 , the primary apertures  307  are configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards a radially inward side  606  of the bulkhead portion  700  of the combustion liner  600 . In an embodiment, the primary apertures  307  in the radially inward section  616  may include a primary aperture angle β 1  configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards the radially inward side  606  of the bulkhead portion  700  of the combustion liner  600   
     In the first section  618 , the primary apertures  307  are configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards a first side  608  of the bulkhead portion  700  of the combustion liner  600 . In an embodiment, the primary apertures  307  in the first section  618  may include a primary aperture angle β 1  configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards the first side  608  of the bulkhead portion  700  of the combustion liner  600 . 
     In the second section  622 , the primary apertures  307  are configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards a second side  612  of the bulkhead portion  700  of the combustion liner  600 . In an embodiment, the primary apertures  307  in the second section  622  may include a primary aperture angle β 1  configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards the second side  612  of the bulkhead portion  700  of the combustion liner  600 . 
     In the center section  624 , the primary apertures  307  are configured to direct the airflow  590  and/or particulate  592  (not shown in  FIG. 3 c   ) towards to central side  615  of the bulkhead portion  700  of the combustion liner  600 . In an embodiment, the primary apertures  307  in the center section  624  may include a primary aperture angle β 1  configured to direct the airflow  590  and/or pollutant particulate  592  (not shown in  FIG. 3 c   ) towards the central side  615  of the bulkhead portion  700  of the combustion liner  600 . 
     It is understood that a combustor of a gas turbine engine is used for illustrative purposes and the embodiments disclosed herein may be applicable to applications other than a combustor of a gas turbine engine. 
     Technical effects of embodiments of the present disclosure include directing impingement airflow within an impingement cavity to reduce airflow speed loss that results in particulate collection with the impingement cavity. 
     The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a non-limiting range of ±8% or 5%, or 2% of a given value. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof. 
     While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.