Abstract:
In accordance with one aspect of the disclosure, a combustor is disclosed. The combustor may include a shell and a liner disposed within the shell. The combustor may further include a grommet at least partially defining a hole communicating through the shell and liner and a cooling channel communicating through the grommet.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This patent application is a US National Stage under 35 U.S.C. §371, claiming priority to International Application No. PCT/US13/021718 filed on Jan. 16, 2013. 
     
    
     FIELD OF THE DISCLOSURE 
       [0002]    The present disclosure generally relates to gas turbine engines and, more specifically, to cooling of combustors in gas turbine engines. 
       BACKGROUND OF THE DISCLOSURE 
       [0003]    A gas turbine engine, typically used as a source of propulsion in aircraft, operates by drawing in ambient air, combusting that air with a fuel, and then forcing the exhaust from the combustion process out of the engine. A fan on a forward end of the engine rotates to draw in ambient air. The air is then compressed by a compressor section having a low-pressure and high-pressure compressor. A portion of the compressed air is used to cool the combustor, while the rest is mixed with a fuel and ignited. 
         [0004]    Typically, an igniter generates an electrical spark to ignite the air-fuel mixture. The products of the combustion, water, CO 2 , NOx, and CO, then travel out of the combustor and exhaust through a turbine. The turbine section, also having a low-pressure and high-pressure turbine, is forced to rotate as the exhaust exits the engine. The turbine section and the compressor section are connected by two concentrically mounted rotating shafts running through the center of the engine. One shaft connects the low-pressure compressor and turbine, while the other shaft connects the high-pressure compressor and turbine. Thus, as the turbine section rotates from the exhaust, the compressor section rotates to bring in and compress new air. Once started, it can therefore be seen that this process is self-sustaining. 
         [0005]    Combustors for gas turbine engines typically have an outer combustor shell and an outer liner, which may be made of a plurality of panels, disposed radially inside the outer combustor shell. Additionally, annular combustors have an inner shell and an inner liner radially outside of the inner shell. The inner and outer liners are separated by and define an annular combustion chamber. Flow cavities are typically provided between each pair of shells and liners. Cooling air is forced through these flow cavities and into the combustion chamber, creating a cooling film on hot surfaces of the liners. 
         [0006]    The remaining portion of the compressed air is used as dilution air to fully burn all of the fuel in the combustion chamber and reduce the temperature of the exhaust. This dilution air is typically injected into a rear section of the combustion chamber through a plurality of holes defined by a plurality of grommets. In prior art designs, engines did not provide any extra cooling for such grommets, and indeed until recent improvements in the design of combustors, extra cooling was often not needed. However, as combustors have advanced, to increase engine power, the temperatures in the combustion chambers have increased. Advanced cooling for the combustor, including for the grommets, is therefore needed. If these areas are not adequately cooled, spallation of the liner, loss of combustor liner material, and cracks or other heat stress related fatigue may occur. 
       SUMMARY OF THE DISCLOSURE 
       [0007]    In accordance with one aspect of the disclosure, a combustor is disclosed. The combustor may include a shell and a grommet at least partially defining a hole communicating through at least the shell, and a cooling channel communicating through the grommet. 
         [0008]    In a refinement, the cooling channel may be oriented perpendicular to a radially inward surface of the grommet with respect to an axis of the combustor extending longitudinally through the combustor. 
         [0009]    In another refinement, the cooling channel may be provided at a non-perpendicular angle to the radially inward surface of the grommet. 
         [0010]    In a further refinement, the cooling channel may communicate through the grommet from a surrounding surface of the grommet oriented perpendicular to the shell of the combustor and facing away from the hole defined y the grommet to the radially inward surface of the grommet. 
         [0011]    In another further refinement, the shell may be engaged with a radially outward surface of the grommet with respect to the axis of the combustor. 
         [0012]    In another refinement, between six to sixteen cooling channels may communicate through the grommet. 
         [0013]    In yet another refinement, each cooling channel may be separated by a distance about equal to three to ten times the diameter of the cooling channels. 
         [0014]    In yet another refinement, the grommet may have a second outward surface with respect to the axis of the combustor engaged with the shell of the combustor. 
         [0015]    In still another refinement, the grommet may be unitary with the shell and may define a hole communicating through the shell. 
         [0016]    In still yet another refinement, the grommet may be separate from the shell and a liner of the combustor, the liner being positioned radially inside the shell with respect to the axis of the combustor, and the grommet being positioned between the shell and liner. 
         [0017]    In accordance with another embodiment, a liner of a combustor is disclosed. The liner may include a liner panel having a hot surface and a grommet defining a hole communicating through the liner panel. The liner may further include a cooling channel communicating through the grommet. 
         [0018]    In a refinement, the grommet may have a radially inward surface and a radially outward surface with respect to an axis extending longitudinally through the combustor, and the cooling channel may extend from the radially outward surface of the grommet to the radially inward surface of the grommet. 
         [0019]    In another refinement, the cooling channel may be at a non-perpendicular angle to the radially inward surface of the grommet. 
         [0020]    In a further refinement, the cooling channel may communicate through the grommet from a surrounding surface of the grommet perpendicular to the radially outward surface of the grommet and facing away from the hole defined by the grommet to the radially inward surface of the grommet. 
         [0021]    In another refinement, between six to sixteen cooling channels may communicate through the grommet. 
         [0022]    In yet another refinement, each cooling channel may be separated by a distance about equal to three to ten times the diameter of the cooling channels. 
         [0023]    In yet another refinement, the grommet may be unitary with the liner panel and the radially inward surface of the grommet is the same surface as the hot surface f the liner panel. 
         [0024]    In still yet another refinement, the grommet may be a dilution hole grommet and may define a dilution hole. 
         [0025]    In accordance with yet another embodiment, a method of cooling a liner of a combustor is disclosed. The method may include providing a grommet with the liner of the combustor, which may at least partially define a hole through the liner panel. The method may further include directing cooling air through a cooling channel communicating through the grommet and cooling the grommet with the cooling air flowing through the cooling channel by transferring heat from the grommet to the cooling air. 
         [0026]    In a refinement, the method may further include blowing a flame in the combustor off the inward surface of the grommet with the cooling air flowing through the cooling channel. 
         [0027]    These and other aspects and features of the present disclosure will be better understood in light of the following detailed description when read in light of the accompanying drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]      FIG. 1  is a cross section of a gas turbine engine constructed in accordance with the present disclosure. 
           [0029]      FIG. 2  is a partial cross section of an annular combustor of a gas turbine engine constructed in accordance with a first embodiment of the present disclosure. 
           [0030]      FIG. 3  is a plan view of a liner panel of a combustor constructed in accordance with a second embodiment of the present disclosure. 
           [0031]      FIG. 4  is a perspective view of the grommet of  FIG. 2 . 
           [0032]      FIG. 5  is a cross section of a liner and a corresponding shell of the combustor of  FIG. 2 . 
           [0033]      FIG. 6  is a cross-sectional view of a liner and shell of a combustor constructed in accordance with a third embodiment of the present disclosure. 
           [0034]      FIG. 7  is a cross section of a liner and shell of a combustor constructed in accordance with a fourth embodiment of the present disclosure. 
           [0035]      FIG. 8  is a cross section of a liner and shell of a combustor constructed in accordance with a fifth embodiment of the present disclosure. 
           [0036]      FIG. 9  is a cross section of a shell of a single walled combustor constructed in accordance with a sixth embodiment of the present disclosure. 
       
    
    
       [0037]    It should be understood that the drawings are not necessarily to scale and that the disclosed embodiments are sometimes illustrated diagrammatically and in partial views. In certain instances, details which are not necessary for an understanding of this disclosure or which render other details difficult to perceive may have been omitted. It should be understood, of course, that this disclosure is not limited to the particular embodiments illustrated herein. 
       DETAILED DESCRIPTION 
       [0038]    Referring now to the drawings, and with specific reference to  FIG. 1 , a gas turbine engine  100  has a fan  101  and a compressor section  102  provided at a front end  106  of the engine  100 . The compressor section  102 , as illustrated, includes a low-pressure compressor  103  and a high-pressure compressor  105 . The low-pressure compressor  103  is connected to a first shaft  108  and the high-pressure compressor  105  is connected to a second shaft  109 . The second shaft  109  is concentrically mounted around the first shaft  108  and both shafts  108 ,  109  extend along and rotate around a central axis  110  extending longitudinally through the engine  20 . When the fan  101  and compressor section  102  rotate, the fan  101  draws ambient air  112  into the engine  100 , and the compressor section  102  compresses the ambient air  110 . The compressed air  114  may be forced through a diffuser  116  to a combustor  118 . At the combustor  118 , the compressed air  114  is split to be used in multiple ways. 
         [0039]    The combustor  118  has a shell  132  and may include a liner  130  mounted to the shell  132 . In the annular combustor illustrated in  FIG. 2 , the combustor  118  has an outer and inner set of shells and liners with respect to the central axis  110  that cooperate to define and are separated by an annular combustion chamber  135 . A combustor axis  119  extends the longitudinally through the combustor  118  equidistant from the outer and inner shells  132 . The shell  132  and associated liner  130  are separated by and define a flow cavity  170  therebetween. Some of the compressed air  114  may pass through a swirler  124  into the combustion chamber  135  as combustion air  122 . The swirler  124  may create turbulence in the combustion air  122  which mixes the combustion air  122  and a fuel  126  entering the combustion chamber  135  by a fuel injector  128 . The air-fuel mixture may then be ignited by an igniter  129  projecting through the liner  130  and shell  132  of the combustor  118 . The combustion products may then be ejected from the combustion chamber  135  as exhaust  136 . As shown in  FIG. 1 , the exhaust  136  passes through a turbine section  138 , having a high-pressure turbine  139  and a low pressure turbine  141 , before exiting the engine  100 . The high-pressure turbine  139  is also connected to the second shaft  109  and the low-pressure turbine  141  is connected to the first shaft  108  such that when the turbine section  138  is rotated by the kinetic energy of the exhaust  136 , the shafts  108  and  109 , and thus the compressor section  102 , are rotated about the central axis  110 . Thereby the process may draw in more ambient air  112  as the exhaust  136  exits the engine  20  and may be self-sustaining once it has begun. 
         [0040]    The compressed air  114  not entering through the swirlers  124  as combustion air  122  may be used as cooling air  144  and dilution air  146 . The cooling air  144  flows through a plurality of impingement holes  172  communicating through the shell  132  into the flow cavity  170  and through a plurality of effusion holes  174  communicating through the liner  130  into the combustion chamber  135 . The dilution air  146 , on the other hand, may enter the combustion chamber  135  at a rear section  148  through at least one dilution hole  150  communicating through the liner  130  and shell  132 . In some embodiments at least one dilution hole  150  communicates through the liner  130  and shell  132  in a forward section  152  of the combustion chamber  135 . The dilution air  146  is burnt in the combustion chamber  135  to complete the combustion process. Additionally, the dilution air  146  may reduce the temperature of the exhaust  136  before the exhaust  136  reaches the turbine section  138 . 
         [0041]    In one embodiment, as illustrated in  FIG. 3 , the liner  130  may be made up of a plurality of panels  154 . Each of the panels  154  may have at least one dilution hole  150  defined by a generally cylindrical dilution hole grommet  156 . The grommet  156  may be unitary with the liner  130  such as the raised platform illustrated in FIGS.  3  and  6 - 8 , or a separate element of the combustor  118  but in engagement with and positioned between the liner  130  and shell  132  of the combustor  118  as illustrated in  FIGS. 2 ,  4  and  5 . 
         [0042]    Speaking now to the embodiment illustrated in  FIGS. 2 ,  5  and  6 , the grommet  156  is a separate element of the combustor engaged with and positioned between the shell  132  and liner  130  of the combustor  118 . The grommet  156  has a radially outward surface  158  with respect to the combustor axis  119  and at least one cooling channel  160  communicating through the grommet  156  from the outward surface  158  to an opposing radially inward surface  159  of the grommet  156  still with respect to the combustor axis  119 . The radially inward surface  159  may be flush with a hot surface  162  of the liner  130 . 
         [0043]    Turning now to embodiments where the grommet  156  is unitary with the liner  130 , such as illustrated in FIGS.  3  and  6 - 8 . The cooling channels  160  communicate through the grommet  156 , or in these embodiments the raised platform that will herein after be referred to simply as the grommet  156 , to the hot surface  162  of the liner  130 . For all purposes herein, the inward surface  159  of a grommet  156  which is separate from the liner  130  is equivalent to the hot surface  162  of the liner  130  when the grommet  156  is unitary with the liner  130 . 
         [0044]    While ten cooling channels  160  are shown in each grommet  156  in  FIG. 3 , in other embodiments between six and sixteen cooling channels  160  may be provided in the grommet  156  to provide various amounts of cooling air  144  to the radially inward surface  159  of the grommet  156  and/or the hot surface  162  of the liner  130 . However, any number of cooling channels  160  may communicate through the grommet  156  to provide any desired amount of cooling air  144  to the radially inward surface  159  and/or hot surface  162 . In another exemplary embodiment, each of the cooling channels  160  are separated by a distance about equal to three to ten times the diameter  164  of the cooling channels  160  to provide an even distribution of the cooling air  144  on the radially inward surface  159  and/or hot surface  162 . In other embodiments, however, the cooling channels  160  may be separated by any desired distance to provide any desired distribution or concentration of cooling air  144  on the radially inward surface  159  and/or hot surface  162 . 
         [0045]    In some embodiments, the grommet  156  may have a second radially outward surface  166  with respect to the combustor axis  119 , which is engaged to an interior surface  167  of the shell  132  still with respect to the combustor axis  119 , as seen in  FIGS. 5-7 . Such engagement of the surfaces  166  and  167  prevent compressed air  114  from passing between the grommet  156  and shell  132  to enter the flow cavity  170 . The first radially outward surface  158  of the grommet  156  may be flush with an exterior surface  168  of the shell  132  with respect to the combustor axis  119 , as illustrated in  FIGS. 5-7 , or may extend further radially outwards than the shell  132 . 
         [0046]    As illustrated in  FIGS. 6 and 7  the cooling channels  160  may be oriented at a non-perpendicular angle to the radially outward surface  158  of the grommet  156 . Specifically, in  FIG. 7  the cooling channels  160  communicate from a generally cylindrical surrounding surface  178  of the grommet to the hot surface  162  of the liner  130 . The surrounding surface  178  is oriented perpendicular to the radially outward surface  159  and faces away from the dilution hole  150  that is defined by the grommet  156 . Such non-perpendicular cooling channels  160  may also be implemented in grommets  156  which are not unitary with the liner  130 . 
         [0047]    In another embodiment, as can be seen in  FIG. 8 , the grommet  156  may only have a first radially outward surface  159  and not have a second radially outward surface  166 . In this embodiment, the first radially outward surface  159  is engaged with the shell  132  such that no compressed air  114  may flow between the shell  132  and the grommet  156 . The cooling channels  160  of this embodiment communicate from the surrounding surface  178  of the grommet  156  to the hot surface  162  of the liner. As stated before, the cooling channels  160  described above may also be implemented in grommets  156  that are not unitary with the liner  130 . 
         [0048]    In combustors  118  which have no liner  130  but only a shell  132 , such as in a can combustor or a single wall annular combustor as illustrated in  FIG. 9  for example, at least one dilution hole  150  may communicate through the shell  132  of the combustor  118  and be defined by the grommet  156  that, in this embodiment, is unitary with the shell  132 . At least one cooling channel  160  communicates through the unitary grommet  156  and shell  132  to provide a path for cooling air  144  to flow into the combustion chamber  135 . 
         [0049]    The cooling air  144  flowing through the cooling channels  160  described above and illustrated in  FIGS. 2-9  cool the grommets  156 , be they unitary with the shell  132 , liner  130 , or separate from both, by transferring heat from the grommet  156  to the cooling air  144 . 
       INDUSTRIAL APPLICABILITY 
       [0050]    From the foregoing, it can be seen that the technology disclosed herein has industrial applicability in a variety of settings such as, but not limited to, cooling dilution hole grommets and the liner around dilution holes (or other holes) in combustors of gas turbine engines. Such engines may be used, for example, in aircraft to generate thrust or in land-based applications to generate power. This improvement over prior art reduces the temperature of the combustor liner around the dilution holes. The reduction in temperature makes the liner less susceptible to damage by heat during engine operations. Such damage may include spallation of the combustor liner, loss of combustor liner material, and cracks or other heat stress related fatigue in the combustor liner. 
         [0051]    While the present disclosure has been in reference to dilution hole grommets, a gas turbine engine, and an aircraft, one skilled in the art will understand that the teachings herein can be used in other applications as well such as, but not limited to, with igniter hole grommets. It is therefore intended that the scope of the invention not be limited by the embodiments presented herein as the best mode for carrying out the invention, but that the invention will include all embodiments falling within the scope of the claims.