Abstract:
One embodiment of the present invention is a unique dome panel for a gas turbine engine combustor. Another embodiment is a unique gas turbine combustor. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and combustion systems and components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Description:
GOVERNMENT RIGHTS 
       [0001]    The present application was made with the United States government support under Contract No. FA8650-07-C-2803, awarded by the United States Air Force. The United States government may have certain rights in the present application. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to combustors and dome panels for gas turbine engines. 
       BACKGROUND 
       [0003]    Gas turbine engine combustors and dome panels for combustors remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
       SUMMARY 
       [0004]    One embodiment of the present invention is a unique dome panel for a gas turbine engine combustor. Another embodiment is a unique gas turbine combustor. Yet another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and combustion systems and components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  schematically illustrates some aspects of a non-limiting example of a gas turbine engine combustor in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  schematically illustrates some aspects of the gas turbine engine combustor of  FIG. 2 . 
           [0009]      FIGS. 4A-4C  illustrate some aspects of a non-limiting example of a dome panel for a combustor of a gas turbine engine in accordance with an embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0010]    For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
         [0011]    Referring to the drawings, and in particular  FIG. 1 , a non-limiting example of some aspects of a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically depicted. In one form, gas turbine engine  10  is an aircraft propulsion power plant. In other embodiments, gas turbine engine  10  may be a land-based or marine engine. In one form, gas turbine engine  10  is a multi-spool turbofan engine. In other embodiments, gas turbine engine  10  may take other forms, and may be, for example, a turboshaft engine, a turbojet engine, a turboprop engine, or a combined cycle engine having a single spool or multiple spools. 
         [0012]    As a turbofan engine, gas turbine engine  10  includes a fan  12 , a bypass duct  14 , a compressor  16 , a diffuser  18 , a combustor  20 , a turbine  22 , a discharge duct  26  and a nozzle system  28 . Bypass duct  14  and compressor  16  are in fluid communication with fan system  12 . Diffuser  18  is in fluid communication with compressor  16 . Combustor  20  is fluidly disposed between compressor  16  and turbine  22 . In one form, combustor  20  includes an annular combustion liner (not shown in  FIG. 1 ) that contains a continuous combustion process. In other embodiments, combustor  20  may take other forms, and may be, for example and without limitation, a can combustor or a canannular combustor. 
         [0013]    Fan  12  includes a fan rotor system  30 . In various embodiments, fan rotor system  30  includes one or more rotors (not shown) that are powered by turbine  22 . Bypass duct  14  is operative to transmit a bypass flow generated by fan system  12  to nozzle  28 . Compressor  16  includes a compressor rotor system  32 . In various embodiments, compressor rotor system  32  includes one or more rotors (not shown) that are powered by turbine  22 . Each compressor rotor includes a plurality of rows compressor blades (not shown) that are alternatingly interspersed with rows of compressor vanes (not shown). Turbine  22  includes a turbine rotor system  34 . In various embodiments, turbine rotor system  34  includes one or more rotors (not shown) operative to drive fan rotor system  30  and compressor rotor system  32 . Each turbine rotor includes a plurality of turbine blades (not shown) that are alternatingly interspersed with rows of turbine vanes (not shown). 
         [0014]    Turbine rotor system  34  is drivingly coupled to compressor rotor system  32  and fan rotor system  30  via a shafting system  36 . In various embodiments, shafting system  36  includes a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed. Turbine  22  is operative to discharge an engine  10  core flow to nozzle  28 . 
         [0015]    In one form, fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  rotate about an engine centerline  48 . In other embodiments, all or parts of fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  may rotate about one or more other axes of rotation in addition to or in place of engine centerline  48 . 
         [0016]    Discharge duct  26  extends between a discharge portion  40  of turbine  22  and engine nozzle  28 . Discharge duct  26  is operative to direct bypass flow and core flow from a bypass duct discharge portion  38  and turbine discharge portion  40 , respectively, into nozzle  28 . In some embodiments, discharge duct  26  may be considered a part of nozzle  28 . Nozzle  28  is in fluid communication with fan system  12  and turbine  22 . Nozzle  28  is operative to receive the bypass flow from., fan system  12  via bypass duct  14 , and to receive the core flow from turbine  22 , and to discharge both as an engine exhaust flow, e.g., a thrust-producing flow. In other embodiments, other nozzle arrangements may be employed, including separate nozzles for each of the core flow and the bypass flow. 
         [0017]    During the operation of gas turbine engine  10 , air is drawn into the inlet of fan  12  and pressurized by fan  12 . Some of the air pressurized by fan  12  is directed into compressor  16  as core flow, and some of the pressurized air is directed into bypass duct  14  as bypass flow, which is discharged into nozzle  28  via discharge duct  26 . Compressor  16  further pressurizes the portion of the air received therein from fan  12 , which is then discharged into diffuser  18 . Diffuser  18  reduces the velocity of the pressurized air, and directs the diffused core airflow into combustor  20 . Fuel is mixed with the pressurized air in combustor  20 , which is then combusted. The hot gases exiting combustor  20  are directed into turbine  22 , which extracts energy in the form of mechanical shaft power sufficient to drive fan  12  and compressor  16  via shafting system  36 . The core flow exiting turbine  22  is directed along an engine tail cone  42  and into discharge duct  26 , along with the bypass flow from bypass duct  14 . Discharge duct  26  is configured to receive the bypass flow and the core flow, and to discharge both into nozzle  28  as an engine exhaust flow, e.g., for providing thrust, such as for aircraft propulsion. 
         [0018]    Referring now to  FIGS. 2 and 3 , combustor  20  is a canted combustor, which is canted at a cant angle  50  relative to engine centerline  48 . Canted combustor  20  includes a plurality of dome panels  52 , a combustion liner  54 , a plurality of fuel nozzles  56 , and a heat shield  58 . Fuel nozzle  56  is not shown in  FIG. 3  for purposes of clarity of illustration. Dome panels  52  are disposed circumferentially around the forward end of combustion liner  54 . In one form, combustion liner  54  is an annular combustion liner. In other embodiments, combustion liner  54  may take other forms. Combustion liner  54  includes an outer combustion liner  60  and an inner combustion liner  62 . Outer combustion liner  60  includes a mating surface  64  configured for engagement with each dome panel  52 . Inner combustion liner  62  includes a mating surface  66  configured for engagement with each dome panel  52 . 
         [0019]    Each dome panel  52  includes a central portion  68 , an upper contact surface  70  and a lower contact surface  72 . Central portion  68  includes an opening  74  configured to receive at least one of a fuel nozzle  56  and a swirler  76 . In other embodiments, more than one opening  74  may be disposed in dome panel  52  for receiving one or more additional fuel nozzles  56  and/or swirlers  76  and/or one or more other components. In one form, swirler  76  is considered a part of fuel nozzle  56 . In other embodiments, swirler  76  may be separate from fuel nozzle  56 . In still other embodiments, combustor  20  may not include a swirler disposed within opening  74 . Opening  74  is canted at cant angle  50 , which orients fuel nozzle  56  at cant angle  50 . In one form, central portion  68  is canted at an angle  78  perpendicular to cant angle  50 . In other embodiments, central portion  68  may be canted at one or more other angles, or may not be canted. 
         [0020]    Referring to  FIGS. 4A-4C , in conjunction with  FIGS. 2 and 3 , upper contact surface  70  extends radially outward from central portion  68  in a radial direction  80  perpendicular to centerline  48  of engine  10 . Lower contact surface  72  extends radially inward from central portion  68  in a radial direction  82  perpendicular to centerline  48  of engine  10 . In one form, central portion  68  is oriented at an angle  84  relative to upper contact surface  70  and lower contact surface  72 . In one form, angle  84  is the same in magnitude as cant angle  50 . In other embodiments, central portion  68  may be oriented differently. 
         [0021]    Upper contact surface  70  is spaced apart from lower contact surface  72  in an axial direction  86  that is parallel to centerline  48  of engine  10 . Upper contact surface  70  is configured for sliding engagement with mating surface  64  of outer combustion liner  60  in directions  80  and  82 . Lower contact surface  72  is configured for sliding engagement with mating surface  66  of inner combustion liner  62  in directions  80  and  82 . Combustion liner  54  and dome panels  52  are thus configured for sliding engagement in directions  80  and  82  perpendicular to centerline  48  of engine  10 . In one form, upper contact surface  70 , lower contact surface  72 , mating surface  64  and mating surface  66  are planar, each having a plane that is perpendicular to centerline of  48  of engine  10 . In other embodiments, one or more of upper contact surface  70 , lower contact surface  72 , mating surface  64  and mating surface  66  may not be planar. The use of at least two planar surfaces, in conjunction with the orientation of at least two planar surfaces in a radial direction permits relative motion between combustion liner  54  and dome panels  52  in radial directions  80  and  82  perpendicular to centerline  48 , which may maintain combustor  20  integrity while undergoing the temperature gradients typically encountered during engine  10  operation. 
         [0022]    In order to aid in mixing fuel and air, and to provide cooling to combustion liner  54 , some embodiments of dome panels  52  include a swirler defined by a plurality of angled openings  88  in central portion  68 . In some embodiments, dome panels  52  also include a deflector  90 , which deflects the air swirled by openings  88  radially outward toward outer combustion liner  60  and inner combustion liner  62  for cooling outer combustion liner  60  and inner combustion liner  62 , as well as along central portion  68  for cooling of dome panels  52 . 
         [0023]    Embodiments of the present invention include a combustor dome panel for a canted combustor of a gas turbine engine, a central portion; an upper contact surface extending radially outward from the central portion in a direction perpendicular to a centerline of the gas turbine engine, wherein the upper contact surface is configured to engage a first mating surface of an outer combustion liner of the canted combustor; and a lower contact surface extending radially inward from the central portion in a direction perpendicular to the centerline of the gas turbine engine, wherein the lower contact surface is configured to engage a second mating surface of an inner combustion liner of the canted combustor. 
         [0024]    In a refinement, the central portion includes an opening configured to receive at least one of a fuel nozzle and a swirler. 
         [0025]    In another refinement, the opening is canted at a cant angle of the canted combustor. 
         [0026]    In yet another refinement, the central portion is canted at an angle perpendicular to a cant angle of the canted combustor. 
         [0027]    In still another refinement, the central portion is oriented at an angle relative to the upper contact surface and the lower contact surface that is the same as a cant angle of the canted combustor. 
         [0028]    In yet still another refinement, the upper contact surface is spaced apart from the lower contact surface in an axial direction parallel to the centerline of the gas turbine engine. 
         [0029]    In a further refinement, the upper contact surface is planar and wherein the lower contact surface is planar. 
         [0030]    Embodiments of the present invention include a canted combustor for a gas turbine engine, comprising: a combustion liner canted at a cant angle relative to a centerline of the gas turbine engine; and a plurality of dome panels configured for mating engagement with the combustion liner, wherein the combustion liner and the plurality of dome panels are configured for sliding engagement in a direction perpendicular to the centerline of the gas turbine engine. 
         [0031]    In a refinement, the sliding engagement is configured to yield relative motion between the combustion liner and the dome panels in a radial direction perpendicular to the centerline of the gas turbine engine. 
         [0032]    In another refinement, at least one dome panel includes an upper contact surface extending radially outward in a direction perpendicular to a centerline of the gas turbine engine; wherein the upper contact surface is configured to engage the combustion liner; wherein the at least one dome panel includes a lower contact surface extending radially inward in a direction perpendicular to the centerline of the gas turbine engine; and wherein the lower contact surface is configured to engage the combustion liner. 
         [0033]    In yet another refinement, the combustion liner includes: an outer combustion liner having a first mating surface configured to engage each dome panel; and an inner combustion liner having a second mating surface also configured to engage each dome panel. 
         [0034]    In still another refinement, the upper contact surface is configured to engage the first mating surface; and wherein the lower contact surface is configured to engage the second mating surface. 
         [0035]    In yet still another refinement, at least one of the upper contact surface and the first mating surface is planar, having a plane perpendicular to the centerline of the gas turbine engine; and wherein at least one of the lower contact surface and the second mating surface is planar, having a plane perpendicular to the centerline of the gas turbine engine. 
         [0036]    In a further refinement, the at least one dome panel includes a canted central portion; wherein the upper contact surface extends radially outward from the canted central portion; and wherein the lower contact surface extends radially inward from the canted central portion. 
         [0037]    In a yet further refinement, the canted central portion is canted at an angle perpendicular to the cant angle of the canted combustor. 
         [0038]    In a still further refinement, the canted central portion is oriented at an angle relative to the upper contact surface and the lower contact surface that is the same as the cant angle of the canted combustor. 
         [0039]    Embodiments of the present invention include a gas turbine engine, comprising: a compressor; a canted combustor in fluid communication with the compressor; and a turbine in fluid communication with the canted combustor, wherein the canted combustor includes a combustion liner and a plurality of dome panels; and wherein the combustion liner and the dome panels are configured for sliding engagement with each other in a direction perpendicular to a centerline of the gas turbine engine. 
         [0040]    In a refinement, at least one dome panel includes: a central portion; an upper contact surface extending radially outward from the central portion in a direction perpendicular to the centerline of the gas turbine engine, wherein the upper contact surface is configured to engage the combustion liner; and a lower contact surface extending radially inward from the central portion in a direction perpendicular to the centerline of the gas turbine engine, wherein the lower contact surface is configured to engage the combustion liner. 
         [0041]    In another refinement, the canted combustor is canted at a cant angle relative to the centerline of the gas turbine engine; and wherein the central portion is canted at the cant angle of the canted combustor relative to the upper contact surface and the lower contact surface. 
         [0042]    In yet another refinement, the canted combustor is canted at a cant angle relative to the centerline of the gas turbine engine; wherein the central portion includes an opening configured to receive at least one of a fuel nozzle and a swirler; and wherein the opening is canted at the cant angle of the canted combustor. 
         [0043]    In still another refinement, the canted combustor includes a combustion liner; wherein the combustion liner includes a first mating surface configured to engage the upper contact surface of each dome panel; wherein the combustion liner includes a second mating surface configured to engage the lower contact surface of each dome panel; and wherein the first mating surface is axially offset from the second mating surface. 
         [0044]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.