Patent Publication Number: US-2019186740-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. 62/607,609 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, and a plurality of cooling holes extending from the second surface to the first surface through the first component; 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; and a plurality of dimples located on the second surface of the second component, the plurality of dimples including: a plurality of primary dimples, wherein each of the plurality of primary dimples is located at an impingement point on the second surface of the second component; and a plurality of secondary dimples, each of the plurality of secondary dimples is located at an impingement flow convergence point. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each impingement point on the second surface of the second component is located opposite a cooling hole of the plurality of cooling holes. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that the second component includes a plurality of secondary cooling holes extending from the second surface to the first surface. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of secondary cooling holes is located between a primary dimple at an impingements point and a secondary dimple at an impingement flow convergence point. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a rounded shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a symmetrical hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a non-symmetrical hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each impingement flow convergence point is located equally between two or more impingement points. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that a height of each of the plurality of the secondary dimple is greater than a height of each of the plurality of primary dimples. 
     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, and a plurality of primary apertures extending from the outer surface to the inner surface through the combustion liner; a heat shield panel interposed between the inner surface the combustion liner and a 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; and a plurality of dimples located on the second surface, the plurality of dimples including: a plurality of primary dimples, wherein each of the plurality of primary dimples is located at an impingement point on the second surface; and a plurality of secondary dimples, each of the plurality of secondary dimples is located at an impingement flow convergence point. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each impingement point on the second surface of the heat shield heat shield panel is located opposite one of the plurality of primary apertures. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the heat shield heat shield panels includes a plurality of secondary apertures extending from the second surface to the first surface. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of secondary apertures is located between a primary dimple at an impingements point and a secondary dimple at an impingement flow convergence point. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a rounded shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a symmetrical hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the plurality of dimples has a non-symmetrical hemispherical shape. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that each impingement flow convergence point is located equally between two or more impingement points. 
     In addition to one or more of the features described above, or as an alternative, further embodiments may include that a height of each of the plurality of the secondary dimple is greater than a height of each of the plurality of primary dimples. 
     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; and 
         FIG. 3 c    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. 
     
    
    
     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 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 that faces a combustion liner inside the combustor. The cooling air may contain particulates, which may build up on the heat shield panels overtime, thus reducing the cooling ability of the cooling air. 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 0.8Mach 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 lbm of fuel being burned divided by lbf 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 . 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  are 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  and 3 c    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 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 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   , particulate  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 particulate  592  impinge upon the second surface  420  of the heat shield panel  400 , the 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 , 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 an 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 dimples  500 , as seen in  FIG. 3 b   . The dimples  600  may be attached to the second surface  410  of the heat shield panel  400  or formed on the second surface  420  of the heat shield panel  400 . Advantageously, the addition of one or more dimples  500  to the second surface of the heat-shield panels  400  may help reduce collection of particulates  592  at the impingement points  594  and the impingement flow convergence points  595  on the second surface  420 . The dimples  500  located at the impingement points  594  may be referred to as primary dimples and the dimples  500  located at the impingement flow convergence points  595  may be referred to as secondary dimples. In an embodiment, the dimples  500  are rounded. In another embodiment, the dimples  500  may each be hemispherical in shape, as seen in  FIG. 3 b   . The hemispherical shape of the dimple  500  may be symmetrical or non-symmetrical. In a symmetrical hemispherically shaped dimple  500 , airflow  590  and particulates  592  would hit a top  510  of the dimple  500  and flow around the dimple  500  in multiple directions. In a non-symmetrical hemispherically shaped dimple  500 , airflow  590  and particulates  592  would hit a top  510  of the shaped dimple  500  and flow around the dimple a selected direction. 
     The airflow  590  and particulate  592  from primary aperture  307  may directed towards a top  510  of the dimple  500 . The top  510  of the dimple  500  may have a rounded shape depending upon the overall shape of the dimple  500 . In an embodiment, the dimples  500  are located such that the secondary apertures  309  are located in between each dimple, as seen in  FIG. 3 b   . The addition of a dimple  500  to the impingement point  594  gently forces airflow  590  and particulate  592  around the dimple  500  and towards the secondary apertures  309 . The rounded shape of the dimple  500  results in less speed loss to the airflow  590  and the particulate  592  than the flat second surface  420  that was seen in  FIG. 3 a   , thus particulate  592  collection at the impingement point  594  on the top  510  of each dimple  500  will be reduced and/or eliminated. 
     As mentioned above, after flowing over the dimple  500  located at an impingement point, some of the airflow  590  and particulate  592  will be directed towards the secondary aperture  309  but then some airflow  590  and particulate will be directed towards a dimple  500  located at the impingement flow convergence point  595 . The rounded shape of the dimple  500  at the impingement flow convergence point  595  will gradually change the direction of the airflow  590  and the particulate  592  so that the airflow  590  and particulate  592  are directed towards the combustion liner  600 , thus the particulate  592  and airflow  590  from a first primary aperture  307  will not have a direct head-on collision when converging with at the 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 . The convergence of particulate  592  and airflow  590  from adjacent primary apertures  307  at the impingent flow convergence point  596  in  FIG. 3 b    may be more of a glancing blow as opposed to a head-on collision, as seen in  FIG. 3 a   , resulting in less speed loss to the airflow  590  and the particulate  592  at the impingement flow convergence point  595  than was seen in  FIG. 3 a   . Thus, advantageously, particulate  592  collection at the impingement flow convergence point  595  will be reduced when utilizing the particulate collection mitigation system  100 , as seen in  FIG. 3   b.    
     Additionally, as shown by the dimples  500  located at the impingement points  594  have a first height H 1  and the dimples  500  located at the impingement flow convergence points  595  may have a second height H 2 . In an embodiment, the first height H 1  may be equivalent to the second height H 2 . In another embodiment, the first height H 1  may be different from the second height H 2 . In a further embodiment, the second height H 2  may be greater than the first height, as seen in  FIG. 3 c   . In another embodiment, the second height H 2  may be greater than or equal to about 50% of the gap height H 3  of the impingement cavity  390 . Advantageously by having a second height H 2  about equivalent to the gap height, particulate may accumulate over time in locations  598 , thus channeling airflow in newly created chambers  599 . 
     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 utilization of dimples on a heat shield panel to maintain speed of particulates and airflow impinging upon the heat shield panel in order to reduce collection of particulate on the heat shield panel. 
     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.