Patent Publication Number: US-2023160574-A1

Title: Gas turbine engine combustor with integral fuel conduit(s)

Description:
BACKGROUND OF THE DISCLOSURE 
     1. Technical Field 
     This disclosure relates generally to a gas turbine engine and, more particularly, to a fuel system for a combustor section of the gas turbine engine. 
     2. Background Information 
     A modern gas turbine engine includes a fuel system for delivering fuel to an array of fuel injectors arranged with a combustor. These fuel injectors inject fuel into a combustion chamber within the combustor for subsequent combustion. Various types and configurations of fuel systems are known in the art. While these known fuel systems have various benefits, there is still room in the art for improvement. There is a need in the art, for example, for a fuel system which can complement operation of another turbine engine component and/or system. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the present disclosure, an assembly is provided for a gas turbine engine. This assembly includes a combustor and a fuel conduit. The combustor includes a combustor wall that forms a peripheral boundary of a combustion chamber within the combustor. The fuel conduit extends along and is formed integral with the combustor wall. The fuel conduit is disposed outside of the combustion chamber. 
     According to another aspect of the present disclosure, another assembly is provided for a gas turbine engine. This assembly includes a combustor and a fuel conduit. The combustor includes an outer sidewall, an inner sidewall, a bulkhead and a combustion chamber. The combustion chamber extends radially within the combustor between the outer sidewall and the inner sidewall. The combustion chamber extends axially within the combustor to the bulkhead. The fuel conduit extends along and is connected to the bulkhead. At least the fuel conduit and the bulkhead are formed together in a monolithic body. 
     According to still another aspect of the present disclosure, another assembly is provided for a gas turbine engine. This assembly includes a combustor, a fuel conduit and a fuel injector. The combustor includes a combustor wall that forms a peripheral boundary of a combustion chamber within the combustor. The fuel conduit extends along and is thermally coupled with the combustor wall. The fuel conduit is arranged on an exterior of the combustor and outside of the combustion chamber. The fuel injector is configured to receive fuel from the fuel conduit and inject the fuel into the combustion chamber. 
     The fuel conduit may extend along and may be connected to the outer sidewall. 
     The fuel conduit may be disposed outside of the combustor. 
     The assembly may also include a turbine engine case and a fuel manifold. The turbine engine case may circumscribe and may be spaced from the combustor. The fuel manifold may extend along and may be connected to the turbine engine case. The fuel manifold may be fluidly coupled with the fuel conduit. 
     The monolithic body may also include the turbine engine case and the fuel manifold. 
     The assembly may also include a fuel injector arranged with the combustor. The fuel injector may be configured to receive fuel from the fuel conduit and inject the fuel into the combustion chamber. 
     The fuel injector may be formed integral with the combustor and the fuel conduit. 
     The assembly may also include a turbine engine case, a plenum and a fuel manifold. The plenum may be formed by and disposed between the turbine engine case and the combustor wall. The fuel manifold may extend along and may be formed integral with the turbine engine case. The fuel manifold may be fluidly coupled with the fuel conduit. 
     The assembly may also include a second fuel conduit extending along and formed integral with the combustor wall. The second fuel conduit may be disposed outside of the combustion chamber. The fuel manifold may be fluidly coupled with the second fuel conduit. 
     A segment of the fuel conduit may project out from the combustor, through the plenum, to the turbine engine case. 
     The combustor may include an inner sidewall, an outer sidewall and a bulkhead extending radially between and connected to the inner sidewall and the outer sidewall. The bulkhead may be configured as or otherwise include the combustor wall. 
     The fuel conduit may also extend along and may be formed integral with the outer sidewall. 
     The combustor may include an inner sidewall, an outer sidewall and a bulkhead extending radially between and connected to the inner sidewall and the outer sidewall. The outer sidewall may be configured as or otherwise include the combustor wall. 
     The assembly may also include a fuel injector arranged with the combustor and fluidly coupled with the fuel conduit. A segment of the fuel conduit may be laterally offset from the fuel injector. 
     The assembly may also include a fuel injector arranged with the combustor and fluidly coupled with the fuel conduit. A segment of the fuel conduit may be circumferentially clocked from the fuel injector about a centerline of the gas turbine engine between five degrees and twenty-five degrees. 
     The assembly may also include a fuel injector arranged with the combustor and fluidly coupled with the fuel conduit. A segment of the fuel conduit may be laterally aligned with the fuel injector. 
     At least a segment of the fuel conduit may follow a straight trajectory as the fuel conduit extends along the combustor wall. 
     At least a segment of the fuel conduit may follow a non-straight trajectory as the fuel conduit extends along the combustor wall. 
     A segment of the fuel conduit may extend along and/or may be spaced from the combustor wall. 
     The assembly may also include a nozzle. The combustor may meet the nozzle at a slip joint between the combustor and the nozzle. 
     The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof. 
     The foregoing features and the operation of the invention will become more apparent in light of the following description and the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a partial side sectional illustration of an assembly for a gas turbine engine. 
         FIG.  2    is a schematic, perspective illustration of a fuel system for the gas turbine engine. 
         FIG.  3    is a partial side sectional illustration of the fuel system configured with a combustor and a diffuser structure. 
         FIG.  3 A  is a side sectional illustration of a slip joint between a combustor and a nozzle. 
         FIG.  3 B  is a side sectional illustration of a length of the fuel conduit selectively spaced from the combustor. 
         FIG.  4    is a partial cutaway illustration of the turbine engine assembly, where the fuel conduit for the fuel system has a first configuration. 
         FIG.  5    is a partial cutaway illustration of the turbine engine assembly, where the fuel conduit has a second configuration. 
         FIG.  6    is a partial, schematic illustration of a gas turbine engine which may include the turbine engine assembly. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    is a partial side sectional illustration of an assembly  10  for a gas turbine engine. This turbine engine assembly  10  includes a (e.g., annular) diffuser structure  12 , a (e.g., annular) combustor  14  and an exit nozzle  16 . The turbine engine assembly  10  also includes a fuel system  18 . 
     The diffuser structure  12  of  FIG.  1    includes one or more cases  20  and  22  of the gas turbine engine. The outer case  20  may be configured as an outer diffuser wall. The inner case  22  may be configured as an inner diffuser wall, an inner turbine wall and/or an exhaust manifold. Each of these turbine engine cases  20 ,  22  extends axially along an axial centerline  24  of the turbine engine assembly  10 , which centerline  24  may also be an axial centerline and/or a rotational axis of the gas turbine engine. Each of the turbine engine cases  20 ,  22  extends circumferentially about (e.g., completely around) the axial centerline  24 . The outer case  20  circumscribes, axially overlaps and is disposed outward of the inner case  22 . The outer case  20  is connected to the inner case  22  at (e.g., on, adjacent or proximate) an (e.g., aft and/or downstream) end  26  of the diffuser structure  12 . The outer case  20  of  FIG.  1    is otherwise spaced radially outward from the inner case  22  to form a diffuser plenum  28  (e.g., an open internal volume, a cavity, etc.) between the turbine engine cases  20  and  22 . 
     The combustor  14  is disposed within the diffuser structure  12  radially between the outer case  20  and the inner case  22 . The combustor  14  includes one or more combustor walls, which combustor walls of  FIG.  1    include a (e.g., tubular) combustor outer sidewall  30 , a (e.g., tubular) combustor inner sidewall  32  and a (e.g., annular) combustor bulkhead  34 ; e.g., a combustor dome. Each of the combustor sidewalls  30 ,  32  extends axially along the axial centerline  24  to the bulkhead  34 . Each of the combustor sidewalls  30 ,  32  extends circumferentially about (e.g., completely around) the axial centerline  24 . The outer sidewall  30  circumscribes, axially overlaps and is radially spaced from the inner sidewall  32 , and the outer case  20  circumscribes, axially overlaps and is radially spaced from the outer sidewall  30 . The inner sidewall  32  circumscribes, axially overlaps and is radially spaced from the inner case  22 . The bulkhead  34  extends radially between and is connected to the outer sidewall  30  and the inner sidewall  32  towards the diffuser structure end  26 . The bulkhead  34  is axially spaced from the outer case  20 . 
     With the foregoing arrangement, the diffuser plenum  28  of  FIG.  1    is formed by and extends radially and/or axially between the diffuser structure  12  and the combustor  14 . More particularly, an (e.g., annular) outer region  36  of the diffuser plenum  28  is formed by and extends radially between and to the outer sidewall  30  and the outer case  20 . An (e.g., annular) inner region  38  of the diffuser plenum  28  is formed by and extends radially between and to the inner sidewall  32  and the inner case  22 . An (e.g., annular) end region  40  of the diffuser plenum  28  is formed by and extends axially between and to the bulkhead  34  and the outer case  20 . 
     The combustor  14  is configured with an internal (e.g., annular) combustion chamber  42 . This combustion chamber  42  extends circumferentially about (e.g., completely around) the axial centerline  24  within the combustor  14 . The combustion chamber  42  extends radially within the combustor  14  between and to the outer sidewall  30  and the inner sidewall  32 . The combustion chamber  42  extends axially within the combustor  14  to the bulkhead  34 . The outer sidewall  30  thereby forms a radial outer peripheral boundary of the combustion chamber  42 . The inner sidewall  32  forms a radial inner peripheral boundary of the combustion chamber  42 . The bulkhead  34  forms an axial end peripheral boundary of the combustion chamber  42 . 
     The exit nozzle  16  is disposed at an outlet of the combustor  14 . The exit nozzle  16  is configured to condition combustion products exiting the combustion chamber  42 . The exit nozzle  16  of  FIG.  1   , for example, includes one or more guide vanes  44  (one visible in  FIG.  1   ) which impart swirl to the combustion products exiting the combustion chamber  42 . The guide vanes  44  are distributed circumferentially about the axial centerline  24  in an annular array. Each of the guide vanes  44  extends radially between and is connected to an inner platform  46  and an outer platform  48 . The inner platform  46  may be aligned with, connected to and/or part of the inner case  22 . The outer platform  48  may be aligned with, connected to and/or part of the outer sidewall  30  and/or another case  50  of the gas turbine engine, which turbine engine case  50  may be configured as an outer turbine case and/or an inner plenum inlet wall. 
     Referring to  FIG.  2   , the fuel system  18  includes a fuel manifold  52 , one or more fuel conduits  54  and one or more fuel injectors  56 . The fuel system  18  also includes at least one feed line  57 ; e.g., an inlet conduit. The fuel manifold  52  is configured to supply fuel received from a fuel source through the feed line  57  to the fuel conduits  54 , which may be fluidly coupled to the fuel manifold  52  in parallel. Each of the fuel conduits  54  is configured to direct the fuel received from the fuel manifold  52  to a respective one of the fuel injectors  56 . Each of the fuel injectors  56  is configured to direct (e.g., inject) the fuel received from the respective fuel conduit  54  into the combustion chamber  42  (see  FIG.  3   ) (or another volume such as a flow tube) for subsequent combustion within the combustion chamber  42 . 
     The fuel manifold  52  of  FIG.  2    extends circumferentially about (e.g., completely around) the axial centerline  24 . The fuel manifold  52  of  FIG.  3    is arranged at an exterior  58  of the diffuser structure  12  proximate the diffuser structure end  26 ; see  FIG.  1   . An entirety of the fuel manifold  52 , for example, may be located outside of the diffuser plenum  28  and, more generally, outside of the diffuser structure  12  and its turbine engine cases  20  and  22 ; see also  FIG.  1   . The fuel manifold  52  is connected to and extends longitudinally (e.g., in a circumferential direction; see  FIG.  2   ) along the outer case  20 . The fuel manifold  52  of  FIG.  3   , for example, is formed integral with the outer case  20 . Herein, the term “integral” may describe an element that is formed monolithically with, non-servable from, as a common part with and/or otherwise configured with another element such that those elements, for example, share at least one common feature and/or structure. The fuel manifold  52  of  FIG.  3   , for example, is configured with the outer case  20  such that a portion  60  of the outer case  20  forms a portion  62  of the fuel manifold  52  and/or the portion  62  of the fuel manifold  52  forms the portion  60  of the outer case  20 . Such an integral configuration may also (e.g., directly) thermally couple the integrally formed elements together to facilitate enhance heat transfer between those elements. 
     The fuel conduits  54  of  FIG.  2    are distributed circumferentially about the axial centerline  24  in an annular array. Each of these fuel conduits  54  extends longitudinally between and is fluidly coupled to the fuel manifold  52  and a respective one of the fuel injectors  56 . Each fuel conduit  54  of  FIG.  3    is arranged at an exterior  64  of the combustor  14 . An entirety of each fuel conduit  54 , for example, may be located outside of the combustion chamber  42  and, more generally, outside of the combustor  14  and its combustor walls  30  and  34 ; see also  FIG.  1   . Each fuel conduit  54  is connected to and extends longitudinally along (e.g., in a radial, an axial and/or a circumferential direction; see also  FIG.  2   ) the combustor  14 . Each fuel conduit  54  of  FIG.  3   , for example, is formed integral with the outer sidewall  30  and the bulkhead  34 . However, an upstream segment  66  of each fuel conduit  54  may project out from the combustor  14  and its bulkhead  34 , axially through the diffuser plenum  28 , to the outer case  20  to meet/couple with the fuel manifold  52 . This upstream segment  66  of the fuel conduit  54  may also be formed integral with (or otherwise connected to) the outer case  20  and/or the fuel manifold  52 . 
     During gas turbine operation, there may be thermal growth differences between the outer case  20  and the combustor  14  and its combustor walls  30 ,  32  and  34  (see also  FIG.  1   ). To reduce thermally induced stresses on the upstream segments  66  (portions of the fuel conduits  54  extending between the combustor  14  and the outer case  20 ), the combustor  14  and, for example, its bulkhead  34  may be rigidly attached to the outer case  20  at (e.g., in, adjacent or proximate) locations of the upstream segments  66 . The combustor  14  of  FIG.  1   , however, may be structurally decoupled from the outer case  20  forward of the upstream segments  66  (see also  FIG.  3   ) to facilitate axially growth of the combustor  14  and its combustor walls  30 ,  32  and  34  along the axial centerline  24  towards the exit nozzle  16  (see  FIG.  1   ) independent of the outer case  20 . In addition or alternatively, referring to  FIG.  3 A , the combustor  14  and its outer sidewall  30  may meet the exit nozzle  16  and its outer platform  48  at a slip joint  67  between the combustor  14  and the exit nozzle  16 . The outer sidewall  30  of  FIG.  3 A , for example, circumscribes (or alternatively may be circumscribed by) and axially overlaps the outer platform  48 . The outer sidewall  30  radially engages (e.g., contacts) and may be axially movable (e.g., slidable) along the outer platform  48 . The slip joint  67  may thereby facilitate relative thermally induced axial movement between the combustor  14  and the exit nozzle  16 . 
     The fuel injectors  56  of  FIG.  2    are distributed circumferentially about the axial centerline  24  in an annular array. Each of these fuel injectors  56  is fluidly coupled with a respective one of the fuel conduits  54 . Each of the fuel injectors  56  of  FIG.  3    is configured with (e.g., arranged with and/or connected to) the combustor  14  and its outer sidewall  30 . Each fuel injector  56  of  FIG.  3   , for example, is connected to the combustor  14  and its outer sidewall  30 , and arranged at a port  68  (e.g., a dilution chute, an opening, a through-hole, a flow guide, etc.) to the combustion chamber  42 . Each fuel injector  56  of  FIG.  3   , for example, is disposed at a distal end of the respective fuel conduit  54 , and formed integral with the outer sidewall  30 . Each fuel injector  56 , for example, may be configured as a nozzle  70  (e.g., an outlet orifice) in an endwall  72  for the respective fuel conduit  54 . Alternatively, one or more of the fuel injectors  56  may be formed discrete from the respective fuel conduit  54 . 
     During operation of the turbine engine system  10  of  FIG.  1   , compressed gas (e.g., air) is directed from a compressor section  74  (e.g., see  FIG.  6   ) through an (e.g., annular) combustor section inlet passage  76  into the diffuser plenum  28 . This compressed gas is diffused within the diffuser plenum  28  and subsequently enters the combustion chamber  42  through the ports  68  (e.g., dilution chutes) and/or one or more other openings into the combustion chamber  42 . Concurrently, the fuel manifold  52  directs fuel into each fuel conduit  54 . Each fuel conduit  54  directs the fuel received from the fuel manifold  52  to the respective fuel injector  56 . Each fuel injector  56  directs (e.g., injects) the fuel received from the respective fuel conduit  54  into the combustion chamber  42  (or another upstream volume) for subsequent mixing with the diffused compressed gas and combustion within the combustion chamber  42 . The combustion products are directed out of the combustor  14  and its combustion chamber  42  through the exit nozzle  16  and to a turbine section  78  (e.g., see  FIG.  6   ). 
     During the combustion process, thermal energy transfers from the combustion products into the combustor walls  30 ,  32  and  34  thereby heating the combustor walls  30 ,  32  and  34 . Some of this thermal energy is transferred into the fuel flowing through the fuel conduits  54 . This transfer of thermal energy heats the fuel; e.g., preheats the fuel for enhanced combustion. The heating of the fuel may be tailored by routing the fuel conduits  54  along the combustor walls  30  and/or  34  to overlap otherwise hot spots and/or cold spots in the combustor walls  30  and/or  34 . To further tailor heating of the fuel, referring to  FIG.  3 B , one or more segments of a fuel conduit  54  may extend along and may be spaced (e.g., via a gap) from the combustor  14  and its the walls  30  and/or  34 . Integrating the fuel conduits  54  with the combustor  14  and its combustor walls  30  and/or  34  as well as the outer case  20  as described above may also simplify manufacture of the turbine engine system  10  where, for example, the turbine engine system  10  is formed using additive manufacturing. 
     In some embodiments, referring to  FIG.  4   , at least a segment or an entirety of one or more or all of the fuel conduits  54  may each be laterally (e.g., circumferentially) aligned with the respective fuel injector  56  connected to that fuel conduit  54 . The fuel conduit  54  and the fuel injector  56  of  FIG.  4   , for example, are circumferentially clocked about the axial centerline  24  at a common circumferential position. In other embodiments, referring to  FIG.  5   , at least a segment or an entirety of one or more or all of the fuel conduits  54  may each be laterally offset from/misaligned with the respective fuel injector  56  connected to that fuel conduit  54 . A segment  80  of the fuel conduit  54  integrated with and extending along the bulkhead  34  and/or a segment  82  of the fuel conduit  54  integrated with and extending along the outer sidewall  30  of  FIG.  5   , for example, is/are circumferentially clocked from the fuel injector  56  about the axial centerline  24  between five degrees(5°) and twenty-five degrees (25°); e.g., between five degrees) (5° and fifteen degrees(15°) or between fifteen degrees(15°) and twenty-five degrees (25°). Of course, in other embodiments, the fuel conduit segment  80 ,  82  may be clocked from the fuel injector  56  about the axial centerline  24  less than five degrees (5°), but more than zero degrees) (0°. In still other embodiments, the fuel conduit segment  80 ,  82  may be clocked from the fuel injector  56  about the axial centerline  24  more than twenty-five degrees (25°). 
     In some embodiments, referring to  FIG.  4   , at least a segment or an entirety of one or more or all of the fuel conduits  54  may each follow a straight longitudinal trajectory  84 A along the combustor  14  when viewed, for example, along the exterior  64  of the combustor  14 ; e.g., a plane parallel with/tangent to the combustor wall  30 ,  34 . The longitudinal trajectory  84 A of the fuel conduit  54  of  FIG.  4   , for example, only includes an axial component and a radial component (e.g., see  FIG.  3   ) as the fuel conduit  54  extends from the fuel manifold  52  to the fuel injector  56 . This may provide a relatively short flowpath from the fuel manifold  52  to the fuel injector  56 , which may provide relatively minimal fuel preheating and/or combustor wall cooling. In other embodiments, referring to  FIG.  5   , at least a segment or an entirety of one or more or all of the fuel conduits  54  may each follow a non-straight (e.g., a curved, a tortuous, a compound, etc.) trajectory  84 B along the combustor  14  when viewed, for example, along the exterior  64  of the combustor  14 ; e.g., a plane parallel with/tangent to the combustor wall  30 ,  34 . The longitudinal trajectory  84 B of the fuel conduit  54  of  FIG.  5   , for example, includes an axial component, a radial component (e.g., see  FIG.  3   ) and a circumferential component as the fuel conduit  54  extends from the fuel manifold  52  to the fuel injector  56 . This may provide a relatively long flowpath from the fuel manifold  52  to the fuel injector  56 , which may provide increased fuel preheating and/or combustor wall cooling. 
     Various components of the turbine engine assembly  10  of  FIG.  1    may be formed together in/as a monolithic body. Herein, the term “monolithic” may describe an apparatus which is formed as a single unitary body. For example, any adjacent two or more or all of the turbine engine components  20 ,  22 ,  30 ,  32 ,  34 ,  44 ,  46 ,  48 ,  50 ,  52 ,  54 ,  56  and/or  57  (see also  FIG.  2   ) (e.g., the entire turbine engine assembly  10 ) may be additively manufactured, cast, machined and/or otherwise formed together as a unitary body. By contrast, a non-monolithic structure includes discretely formed bodies that are subsequently mechanically fastened and/or otherwise removably attached together. 
       FIG.  6    schematically illustrates a single spool, radial-flow turbojet gas turbine engine  86  with which the turbine engine assembly  10  may be included. This gas turbine engine  86  may be configured for propelling an unmanned aerial vehicle (UAV), a drone, or any other manned or unmanned aircraft or self-propelled projectile. In the specific embodiment of  FIG.  6   , the gas turbine engine  86  includes an upstream inlet  88 , the (e.g., radial) compressor section  74 , a combustor section  90  with the combustor  14  and the combustion chamber  42 , the (e.g., radial) turbine section  78  and a downstream exhaust  92  fluidly coupled in series. A compressor rotor  94  in the compressor section  74  is coupled with a turbine rotor  96  in the turbine section  78  by a shaft  98 , which shaft  98  may rotate about the axial centerline  24 . 
     The gas turbine engine  86  is described above as a single spool, radial-flow turbojet gas turbine engine for ease of description. The present disclosure, however, is not limited to such an exemplary gas turbine engine. The gas turbine engine  86 , for example, may alternatively be configured as an axial flow gas turbine engine. The gas turbine engine  86  may be configured as a direct drive gas turbine engine. The gas turbine engine  86  may alternatively include a gear train that connects one or more rotors together such that the rotors rotate at different speeds. The gas turbine engine  86  may be configured with a single spool (e.g., see  FIG.  6   ), two spools, or with more than two spools. The gas turbine engine  86  may be configured as a turbofan engine, a turbojet engine, a propfan engine, a pusher fan engine or any other type of turbine engine. In addition, while the gas turbine engine  86  is described above with an exemplary reverser flow annular combustor, the gas turbine engine  86  may also or alternatively include any other type/configuration of annular, tubular (e.g., CAN), axial flow and/or reverser flow combustor. The present disclosure therefore is not limited to any particular types or configurations of gas turbine engines. 
     While various embodiments of the present disclosure have been described, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.