Patent Publication Number: US-2023160323-A1

Title: Sectioned engine structure for a gas turbine engine

Description:
BACKGROUND OF THE DISCLOSURE 
     1. Technical Field 
     This disclosure relates generally to a gas turbine engine and, more particularly, to a stationary structure for the gas turbine engine. 
     2. Background Information 
     A gas turbine engine includes a stationary engine structure for housing and/or supporting internal rotating components of the gas turbine engine. A typical stationary engine structure includes a plurality of tubular axial case segments. These tubular axial case segments are arranged sequentially along an axial centerline of the gas turbine engine and axially connected together by flange connections. While such a stationary engine structure has various benefits, there is still room in the art for improvement. There is a need in the art therefore for an improved stationary engine structure as well as methods for manufacturing and assembling stationary engine structure components. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the present disclosure, an assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a stationary engine structure. The stationary engine structure includes a diffuser, a combustor, an engine case and a plenum. The combustor is disposed within the plenum. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. A first section of the stationary engine structure is formed as a first monolithic body. The first section includes the diffuser and the combustor. A second section of the stationary structure is formed as a second monolithic body. The second section is configured as or otherwise includes the engine case. 
     According to another aspect of the present disclosure, another assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a diffuser, a combustor, a duct wall and an engine wall. The diffuser includes an inner diffuser wall and an outer diffuser wall. The combustor includes an inner combustor wall, an outer combustor wall and a bulkhead extending between and connected to the inner combustor wall and the outer combustor wall. The engine wall includes a side wall and an end wall. The side wall projects out from the end wall to the outer diffuser wall. The side wall is brazed to the outer diffuser wall. The end wall projects out from the side wall to the duct wall. The end wall is brazed to the duct wall. 
     According to still another aspect of the present disclosure, a manufacturing method is provided. During this method, an engine structure preform is formed. The engine structure preform includes a first section preform and a second section preform formed integral with the first section preform. The first section preform includes a diffuser and a combustor. The second section preform is configured as or otherwise includes an engine case. The first section preform is separated from the second section preform to respectively provide a first section and a second section that is discrete from the first section. The second section is attached to the first section to provide an engine structure. The combustor is disposed within a plenum of the engine structure. The engine case forms a peripheral boundary of the plenum. A gas path extends sequentially through the diffuser, the plenum and the combustor. 
     The forming may include additive manufacturing the engine structure preform. In addition or alternatively, the attaching may include bonding the second section to the first section. 
     The assembly may include a monolithic body including the diffuser, the combustor and the duct wall. 
     The diffuser may include a first wall, a second wall and a plurality of vanes. Each of the vanes may be within the gas path and may extend between the first wall and the second wall. 
     The combustor may be configured as a reverse-flow combustor. 
     The stationary engine structure may also include a turbine nozzle downstream of the combustor along the gas path. The first section may include the turbine nozzle. 
     The diffuser and the turbine nozzle may share a common wall. 
     The turbine nozzle may include a first wall, a second wall and a plurality of vanes. Each of the vanes may be within the gas path and may extend between the first wall and the second wall. 
     The assembly may also include a turbine rotor. The stationary engine structure may also include a turbine case housing the turbine rotor and forming a peripheral boundary of the gas path. The first section may also include the turbine case. 
     The stationary engine structure may also include an exhaust duct forming a peripheral boundary of the gas path. The first section may also include the exhaust duct. 
     The second section may circumscribe the first section. 
     The second section may be bonded to the first section through a butt joint. 
     The second section may be bonded to the first section through a splice joint. 
     The engine case may extend axially along and circumferentially about an axis. The engine case may include a side wall and an end wall. The side wall may project axially out from the end wall to an axial end of the engine case. The engine case may be attached to the first section at the axial end. The end wall may project radially in from the side wall to a radial end of the engine case. The engine case may be attached to the first section at the radial end. 
     The stationary engine structure may also include a fuel conduit and a nozzle. The nozzle may be configured to receive fuel from the fuel conduit and inject the fuel into a volume of the combustor. The second section may also include the fuel conduit and/or the nozzle. 
     The stationary structure may also include a fuel manifold outside of the engine case. The fuel manifold may be configured to supply the fuel to the fuel conduit. The second section may also include the fuel manifold. 
     At least a portion of the fuel conduit may project out from the engine case into the plenum towards the fuel injector. 
     The stationary engine structure may also include an inlet section and a compressor case. The compressor case may form a peripheral boundary of the gas path between the inlet section and the diffuser. A third section of the stationary structure may include the inlet section and the compressor case. The third section may be attached to the first section. 
     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 schematic side sectional illustration of a gas turbine engine. 
         FIG.  2    is an aft end view illustration of a stationary engine structure for the gas turbine engine. 
         FIG.  3    is a schematic side sectional illustration of a central portion of the gas turbine engine. 
         FIG.  4    is a schematic perspective illustration of a portion of a fuel delivery system. 
         FIG.  5    is a partial side sectional illustration of an outer section of the stationary engine structure bonded to an inner section of the stationary engine structure. 
         FIG.  6    is a schematic illustration of a manufacturing system. 
         FIG.  7    is a flow diagram of a method for forming a gas turbine engine. 
         FIGS.  8 A and  8 B  are schematic illustrations of various preforms of the stationary engine structure. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    is a side sectional illustration of a gas turbine engine  20 . The gas turbine engine  20  of  FIG.  1    is configured as a single spool, radial-flow turbojet turbine engine. This gas turbine engine  20  is configured for propelling an aircraft such as, but not limited to, an unmanned aerial vehicle (UAV), a drone or any other manned or unmanned aircraft or self-propelled projectile. The present disclosure, however, is not limited to such an exemplary turbojet turbine engine configuration nor to an aircraft propulsion system application. For example, the gas turbine engine  20  may alternatively be configured as an auxiliary power unit (APU) or an industrial gas turbine engine. 
     The gas turbine engine  20  of  FIG.  1    extends axially along an axial centerline  22  between a forward, upstream engine inlet  24  and an aft, downstream engine exhaust  26 . This axial centerline  22  may also be a rotational axis for various components within the gas turbine engine  20 . 
     The gas turbine engine  20  includes a compressor section  28 , a combustor section  30  and a turbine section  32 . The gas turbine engine  20  also includes a stationary engine structure  34 . This stationary engine structure  34  houses the compressor section  28 , the combustor section  30  and the turbine section  32 . The stationary engine structure  34  of  FIG.  1    also forms an inlet section  36  and an exhaust section  38  for the gas turbine engine  20 , where the inlet section  36  forms the engine inlet  24  and the exhaust section  38  forms the engine exhaust  26 . 
     The engine sections  36 ,  28 ,  30 ,  32  and  38  are arranged sequentially along a core gas path  40  that extends through the gas turbine engine  20  from the engine inlet  24  to the engine exhaust  26 . Each of the engine sections  28  and  32  includes a respective rotor  42 ,  44 . Each of these rotors  42 ,  44  includes a plurality of rotor blades arranged circumferentially around and connected to at least one respective rotor disk. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s). 
     The compressor rotor  42  may be configured as a radial flow rotor. The turbine rotor  44  may also or alternatively be configured as a radial flow rotor. The compressor rotor  42  is connected to the turbine rotor  44  through an engine shaft  46 . This shaft  46  is rotatably supported by the stationary engine structure  34  through a plurality of bearings  48 A and  48 B (generally referred to as  48 ); e.g., rolling element bearings, journal bearings, etc. 
     The combustor section  30  includes an annular combustor  50  with an annular combustion chamber  52 . The combustor  50  of  FIG.  1    is configured as a reverse flow combustor. Inlets ports  54  (e.g., dilution chutes) into the combustion chamber  52 , for example, may be arranged at (e.g., on, adjacent or proximate) and/or towards an aft bulkhead wall  56  (e.g., bulkhead, dome, etc.) of the combustor  50 . An outlet from the combustor  50  may be arranged axially aft of an inlet to the turbine section  32 . The combustor  50  may also be arranged radially outboard of and/or axially overlap at least a (e.g., aft) portion of the turbine section  32 . With this arrangement, the core gas path  40  of  FIG.  1    reverses its directions (e.g., from a forward-to-aft direction to an aft-to-forward direction) a first time as the gas path  40  extends from a combustor plenum  58  surrounding the combustor  50  into the combustion chamber  52 . The core gas path  40  of  FIG.  1    then reverses its direction (e.g., from the aft-to-forward direction to the forward-to-aft direction) a second time as the gas path  40  extends from the combustion chamber  52  into the turbine section  32 . 
     During operation, air enters the gas turbine engine  20  through the inlet section  36  and its engine inlet  24 . The inlet section  36  directs this air from the engine inlet  24  into the core gas path  40  and the compressor section  28 . The engine inlet  24  of  FIG.  1    thereby forms a forward, upstream inlet to the core gas path  40  and the compressor section  28 . The air within the core gas path  40  may be referred to as core air. 
     The core air is compressed by the compressor rotor  42  and directed through an annular diffuser  60  and the plenum  58  into the combustion chamber  52 . Fuel is injected and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited within the combustion chamber  52 , and combustion products thereof flow through the turbine section  32  and cause the turbine rotor  44  to rotate. This rotation of the turbine rotor  44  drives rotation of the compressor rotor  42  and, thus, compression of the air received from the engine inlet  24 . The exhaust section  38  receives the combustion products from the turbine section  32 . The exhaust section  38  directs the received combustion products out of the gas turbine engine  20  to provide forward engine thrust. 
     The stationary engine structure  34  of  FIG.  1    may include some or all stationary engine components included in the gas turbine engine  20 . Herein, the term stationary may describe a component that does not rotate with the rotating assembly (e.g., an assembly of the rotors  42  and  44  and the shaft  46 ) during gas turbine engine operation. A stationary component, for example, may refer to any component that remains stationary during gas turbine engine operation such as, but not limited to, a wall, a liner, a strut, a fixed vane, a fuel nozzle, a conduit, etc. 
     The stationary engine structure  34  of  FIG.  1    is configured as a generally tubular structure. The stationary engine structure  34 , for example, extends axially along the axial centerline  22  from the inlet section  36  to the engine section  38 . The stationary engine structure  34  extends circumferentially about (e.g., completely around) the axial centerline  22  such that the stationary engine structure  34  has, for example, a full hoop geometry; see also  FIG.  2   . 
     The stationary engine structure  34  includes one or more case walls. The stationary engine structure  34  of  FIG.  3   , for example, includes an outer compressor wall  62 , an outer diffuser wall  64  of the diffuser  60 , an inner diffuser wall  66  of the diffuser  60 , a plenum side (e.g., outer) wall  68 , a plenum end wall  70 , an outer combustor wall  72  of the combustor  50 , an inner combustor wall  74  of the combustor  50 , the bulkhead wall  56  of the combustor  50 , an inner turbine wall  76  of a turbine duct  78 , and an exhaust wall  80  of an exhaust duct  82  (see also  FIG.  1   ). At least a portion or an entirety of each of the case walls  62 ,  64 ,  66 ,  68 ,  72 ,  74 ,  76  and/or  80  of  FIG.  3   , for example, is generally tubular. At least a portion or an entirety of each of the case walls  56  and/or  70  of  FIG.  3    is generally annular. 
     The compressor wall  62  extends axially along the axial centerline  22  between and is connected to the inlet section  36  and the outer diffuser wall  64 . The compressor wall  62  of  FIG.  3    circumscribes, axially overlaps and thereby houses the compressor rotor  42 . 
     The outer diffuser wall  64  extends axially along the axial centerline  22  between and is connected to the compressor wall  62  and the plenum side wall  68 . The outer diffuser wall  64  is spaced radially outboard from, axially overlaps and circumscribes the inner diffuser wall  66 . The outer diffuser wall  64  of  FIG.  3    thereby forms an outer peripheral boundary of the core gas path  40  through the diffuser  60 . 
     The inner diffuser wall  66  may be connected to outer combustor wall  72 . The inner diffuser wall  66  of  FIG.  3   , for example, projects axially out from the outer combustor wall  72  and extends axially towards (e.g., to) an aft, downstream end of an inner platform of the compressor rotor  42 . This inner diffuser wall  66  forms an inner peripheral boundary of the core gas path  40  within the diffuser  60 . The inner diffuser wall  66  may also be configured as an outer turbine wall. The inner diffuser wall  66  of  FIG.  3   , for example, may also form an outer peripheral boundary of the core gas path  40  within a (e.g., upstream) portion of the turbine section  32 . The inner diffuser wall  66  of  FIG.  3    circumscribes, axially overlaps and may thereby house a (e.g., upstream) portion of the turbine rotor  44 . 
     The plenum side wall  68  extends axially along the axial centerline  22  between and is connected to the outer diffuser wall  64  and the plenum end wall  70 . The plenum side wall  68  of  FIG.  3    circumscribes, axially overlaps and thereby houses the combustor  50  and its outer combustor wall  72 . The plenum side wall  68  is radially spaced outward from the combustor  50  and its outer combustor wall  72 . The plenum side wall  68  forms an outer peripheral boundary of the plenum  58 . 
     The plenum end wall  70  extends radially (and axially along the axial centerline  22 ) between and is connected to the plenum side wall  68  and the exhaust wall  80 . The plenum end wall  70  is axially spaced from the combustor  50  and its bulkhead wall  56 . The plenum end wall  70  forms an axial end peripheral boundary of the plenum  58 . 
     The outer combustor wall  72  extends axially along the axial centerline  22  between and may be connected to the bulkhead wall  56  and the inner diffuser wall  66 . More particularly, the outer combustor wall  72  extends axially to and may be connected to an outer platform  84  of a turbine nozzle  86 ; e.g., an exit nozzle from the combustion chamber  52 . This nozzle outer platform  84  of  FIG.  3    is configured as part of the inner diffuser wall  66 ; however, the walls  66  and  84  may be discrete from one another in alternative embodiments. 
     The inner combustor wall  74  is connected to the bulkhead wall  56 . This inner combustor wall  74  projects axially along the axial centerline  22  out from the bulkhead wall  56  towards the turbine nozzle  86  and its inner platform  88 . This nozzle inner platform  88  of  FIG.  3    is configured as part of the inner turbine wall  76 ; however, the walls  76  and  88  may be discrete from one another in alternative embodiments. 
     The bulkhead wall  56  extends radially between the outer combustor wall  72  and the inner combustor wall  74 . The bulkhead wall  56  is connected to an aft end portion of the outer combustor wall  72  and an aft end portion of the inner combustor wall  74 . With this arrangement, the combustor case walls  56 ,  72  and  74  collectively form peripheral boundaries of the combustion chamber  52  within the combustor  50 . 
     The inner turbine wall  76  may be wrapped around a downstream end portion of the inner combustor wall  74 . An upstream portion of the inner turbine wall  76  of  FIG.  3    (e.g., the inner platform  88 ), for example, circumscribes and axially overlaps the downstream end portion of the inner combustor wall  74 . This upstream portion extends axially along the axial centerline  22  (in the aft-to-forward direction) to a turning portion of the inner turbine wall  76 . A downstream portion of the inner turbine wall  76  projects axially (in the forward-to-aft direction) away from the inner turbine wall turning portion to the exhaust wall  80 . The inner turbine wall  76  is circumscribed and axially overlapped by the combustor  50  and its inner combustor wall  74 . The inner turbine wall  76  is also spaced radially inboard from the combustor  50  and its inner combustor wall  74 . The inner turbine wall  76  of  FIG.  3    forms an inner peripheral boundary of the plenum  58 , where the combustor  50  is disposed within and is substantially surrounded by the plenum  58 . The inner turbine wall  76  forms an outer peripheral boundary of the core gas path  40  within a (e.g., downstream) portion of the turbine section  32 . The inner turbine wall  76  of  FIG.  3    also circumscribes, axially overlaps and thereby houses a (e.g., downstream) portion of the turbine rotor  44 . 
     The exhaust wall  80  is connected to the inner turbine wall  76 . The exhaust wall  80  of  FIG.  1    projects axially out from the inner turbine wall  76  to the aft engine exhaust  26 . 
     The stationary engine structure  34  may include one or more internal support structures with one or more support members. Examples of the support members include, but are not limited to, struts, structural guide vanes, bearing supports, bearing compartment walls, etc. The stationary engine structure  34  of  FIG.  3   , for example, includes a forward support structure  90  to support the forward bearing  48 A and an aft support structure  92  to support the aft bearing  48 B. The stationary engine structure  34  of  FIG.  3    also includes an inlet nozzle  94 , a diffuser nozzle  96  and the turbine nozzle  86 . 
     The inlet nozzle  94  may be configured to condition the core air entering the compressor section  28 . The inlet nozzle  94  of  FIG.  3   , for example, includes one or more inlet guide vanes  98  configured to impart swirl to the core air. These inlet guide vanes  98  are arranged circumferentially about the axial centerline  22  in an annular array. Each of the inlet guide vanes  98  extends radially across the gas path  40 . 
     The diffuser nozzle  96  may be configured to condition the core air leaving the compressor section  28  and entering the plenum  58 . The diffuser nozzle  96  of  FIG.  3   , for example, includes one or more diffuser guide vanes  100  configured to impart swirl to the core air. These diffuser guide vanes  100  are arranged circumferentially about the axial centerline  22  in an annular array. Each of the diffuser guide vanes  100  extends radially across the gas path  40 . More particularly, each of the diffuser guide vanes  100  extends radially between and is connected to the inner diffuser wall  66  and the outer diffuser wall  64 . 
     The turbine nozzle  86  may be configured to condition the combustion products exiting the combustor  50  and its combustion chamber  52 . The turbine nozzle  86  of  FIG.  3   , for example, includes one or more turbine guide vanes  102  configured to impart swirl to the combustion products. These turbine guide vanes  102  are arranged circumferentially about the axial centerline  22  in an annular array. Each of the turbine guide vanes  102  extends radially across the gas path  40 . More particularly, each of the turbine guide vanes  102  extends radially between and is connected to the turbine nozzle outer and inner platforms  84  and  88 . 
     Referring to  FIG.  4   , the stationary engine structure  34  may also include one or more components of a fuel delivery system. These fuel delivery system components may include a fuel feed line  104  (e.g., an inlet conduit), a fuel manifold  106 , one or more fuel conduits  108  and one or more fuel injectors  110 . The fuel manifold  106  is configured to supply fuel received from a fuel source (e.g., a fuel reservoir, a fuel pump, etc.) through the fuel feed line  104  to the fuel conduits  108 , which fuel conduits  108  may be fluidly coupled to the fuel manifold  106  in parallel. Referring to  FIG.  5   , each of the fuel conduits  108  is configured to direct the fuel received from the fuel manifold  106  to a respective one of the fuel injectors  110 . Each of the fuel injectors  110  is configured to direct (e.g., inject) the fuel received from the respective fuel conduit  108  into a respective volume of the combustor  50  for subsequent combustion within the combustion chamber  52 . The combustor volume may be a respective one of the ports  54  in a sidewall of the combustor  50  (e.g., the outer combustor wall  72 ), where the port  54  may be configured as a dilution chute, a flow guide, an orifice or any other opening and/or passage through the combustor sidewall and into the combustion chamber  52 . 
     The fuel manifold  106  of  FIG.  4    extends circumferentially about (e.g., completely around) the axial centerline  22 . Referring to  FIG.  5   , the fuel manifold  106  is arranged at an exterior of an outer engine case  112  (e.g., a plenum case, a combustor section case, etc.) of the stationary engine structure  34 , which outer engine case  112  may include the plenum side wall  68  and the plenum end wall  70 . An entirety of the fuel manifold  106 , for example, may be located outside of the plenum  58  and, more generally, outside of the engine case  112  and its plenum end wall  70 . The fuel manifold  106  is connected to and extends longitudinally (e.g., in a circumferential direction; see  FIG.  4   ) along the engine case  112 . 
     The fuel conduits  108  of  FIG.  4    are distributed circumferentially about the axial centerline  22  in an annular array. Each of these fuel conduits  108  extends longitudinally between and is fluidly coupled to the fuel manifold  106  and a respective one of the fuel injectors  110 . Referring to  FIG.  5   , each fuel conduit  108  is arranged at an interior of the engine case  112 . Each fuel conduit  108  of  FIG.  5   , for example, is connected to the engine case  112  and its plenum end wall  70 . Each fuel conduit  108  projects (e.g., axially) out from the engine case  112  and its plenum end wall  70  into the plenum  58  towards (e.g., to) the respective fuel injector  110 . A first (e.g., upstream) segment  114  of each fuel conduit  108  of  FIG.  5    extends axially along and may be integral with the engine case  112  and one or more of its walls  68  and/or  70 ; e.g., the plenum end wall  70 . A second (e.g., downstream) segment  116  of each fuel conduit  108  of  FIG.  5    projects axially out from the conduit first segment  114 , away from the engine case  112 , towards (e.g., to) the respective fuel injector  110 . This conduit second segment  116  is radially spaced (e.g., separated) from the engine case  112  and one or more of its walls  68  and/or  70 ; e.g., the plenum side wall  68 . A distal end of each fuel conduit  108  and/or each fuel injector  110 , however, may be structurally supported within the plenum  58  by a support  118 ; e.g., a strut, a vane, a post, a beam, etc. This support  118  is configured to maintain a position of the fuel injector  110  during gas turbine engine operation. Without the support  118 , for example, thermal expansion and/or contraction of the fuel conduit  108  may lead to (e.g., axial and/or circumferential) misalignment of the respective fuel injector  110  and port  54 ; e.g., dilution chute. Of course, in other embodiments, an entirety of one or more or all of the fuel conduits  108  may each be formed integral with the engine case  112  to obviate the need for the respective support  118 . 
     The fuel injectors  110  of  FIG.  4    are distributed circumferentially about the axial centerline  22  in an annular array. Each of these fuel injectors  110  is fluidly coupled with a respective one of the fuel conduits  108 . Referring to  FIG.  5   , each of the fuel injectors  110  is (e.g., circumferentially and/or axially) aligned with and located next to the respective port  54 ; e.g., dilution chute. Each fuel injector  110  is disposed at a distal end of the respective fuel conduit  108 , and is formed integral with the respective fuel conduit  108  and support  118 . Each fuel injector  110 , for example, may be configured as a nozzle (e.g., an outlet orifice) in a sidewall of the respective fuel conduit  108 . Alternatively, one or more of the fuel injectors  110  may be formed discrete from the respective fuel conduit  108 . 
     The stationary engine structure  34  of  FIG.  3    is configured from a plurality of discrete sections of the engine structure  34 ; e.g., engine sub-structures. The stationary engine structure  34  of  FIG.  3   , for example, at least or only includes an upstream section  120 , an inner downstream section  121  and an outer downstream section  122 . The upstream section  120  may include an entirety (or at least a portion) of any one or more or all of the engine structure elements  36 ,  62 ,  90 ,  94  and  98 . The inner downstream section  121  may include an entirety (or at least a portion) of any one or more or all of the engine structure elements  50 ,  56 ,  64 ,  66 ,  72 ,  74 ,  76 ,  78 ,  80 ,  82 ,  84 ,  86 ,  88 ,  92 ,  96 ,  100 ,  102 ; see also  FIG.  1   . The outer downstream section  122  may include an entirety (or at least a portion) of any one or more or all of the engine structure elements  68 ,  70 ,  104  (see  FIGS.  2  and  4   ),  106 ,  108 ,  110 ,  112 ,  114 ,  116 ,  118  and  134 . With such an arrangement, the stationary engine structure  34  may be configured from (include) relatively few discrete parts (e.g., discretely formed bodies) while still facilitating inspection and/or finishing (e.g., machining, coating, etc.) interior surfaces of the stationary engine structure  34 . For example, forming the downstream sections  121  and  122  discrete from one another facilitates inspection and/or finishing of the fuel injectors  110  and their orifices, inspection and/or finishing of the ports  54  (e.g., dilution chutes), etc. 
     Each of the engine structure sections  120 - 122  may be formed as a monolithic body. Herein, the term monolithic may described an apparatus which is formed as a single unitary body. Each engine structure section  120 ,  121 ,  122 , for example, may be additively manufactured, cast, machined and/or otherwise formed as an integral, unitary body. By contrast, a non-monolithic body may include parts that are discretely formed from one another, where those parts are subsequently mechanically fastened and/or otherwise attached to one another. 
     The upstream section  120  is mated with and connected to a forward, upstream end of the inner downstream section  121 . The upstream section  120  of  FIG.  3   , for example, is attached to the inner downstream section  121  via at least one mechanical joint  124 ; e.g., a bolted flange connection. However, in other embodiments, the upstream section  120  may also or alternatively be attached to the inner downstream section  121  via at least one bonded joint; e.g., a brazed connection, a welded connection, etc. 
     The outer downstream section  122  is mated with and connected to the inner downstream section  121 . The inner downstream section  121  of  FIG.  5   , for example, is received within a bore of the outer downstream section  122 , where the outer downstream section  122  and its engine case  112  circumscribe and axially overlap the combustor  50  and the exhaust duct  82 . The outer downstream section  122  is connected to the inner downstream section  121  at a forward axial end  126  of the engine case  112  and a radial inner end  128  of the engine case  112 . The plenum side wall  68  of  FIG.  5   , for example, is brazed, welded and/or otherwise bonded to the outer diffuser wall  64  at an (e.g., annular) axial interface  130 ; e.g., a butt joint. An annular end (e.g., edge) of the plenum side wall  68  of  FIG.  5   , for example, is aligned with and axially engages (e.g., via bonding material such as braze material) an annular end (e.g., edge) of the outer diffuser wall  64 . The plenum end wall  70  of  FIG.  5    is brazed, welded and/or otherwise bonded to the exhaust duct  82  and its exhaust wall  80  (and/or the turbine duct  78  and its inner turbine wall  76 ) at a (e.g., tubular) radial interface  132 ; e.g., a splice joint such as an overlap joint. The engine case  112  of  FIG.  5   , for example, includes a tubular flange  134  that projects axially (e.g., in an aft, downstream direction) out from the plenum end wall  70  at the radial inner end  128  to an axial distal end. An inner surface of this tubular flange  134  circumscribes, axially overlaps and radially engages (e.g., via bonding material such as braze material) an outer surface of the exhaust duct  82  and its exhaust wall  80  (and/or the turbine duct  78  and its inner turbine wall  76 ). 
     The axial interface  130  and/or the radial interface  132  may be positioned relatively far from one or more of the fuel delivery system components. The axial interface  130  of  FIG.  5   , for example, is located at (e.g., on, adjacent or proximate) the diffuser  60  and an outlet from the combustor  50 . The radial interface  132  of  FIG.  5    is located at (e.g., on, adjacent or proximate) the exhaust duct  82  (and/or the turbine duct  78 ) and a radial inner side of the combustor  50 . With this arrangement, heat transferred into the outer downstream section  122  and its engine case  112  during bonding of the engine structure sections  121  and  122  may significantly dissipate prior to reaching the fuel delivery system components  108  and/or  110 . The fuel delivery system components  108  and/or  110  may therefore be subject to very little or no thermal distortion during the bonding of the engine structure sections  121  and  122 . The fuel injectors  110  may thereby remain properly positioned relative to (e.g., aligned with) the respective ports  54 . By contrast, if the interfaces  130  and  132  are positioned too close to (e.g., next to) the fuel delivery system components  108  and/or  110 , the fuel delivery system components  108  and/or  110  may be subject to thermal distortion during bonding and the fuel injectors  110  may become misaligned with (e.g., offset from) the respective ports  54 . 
       FIG.  6    illustrates a system  136  for manufacturing a stationary engine structure such as, but not limited to, the stationary engine structure  34  of  FIG.  1   . This manufacturing system  136  (e.g., an additive manufacturing system) includes a build chamber  138  defining a build space  140  for manufacturing the stationary engine structure  34  (or one or more of its components). The manufacturing system  136  also includes an additive manufacturing apparatus  142 ; e.g., a laser powder bed fusion (LPBF) apparatus. 
     The additive manufacturing apparatus  142  is configured to build the stationary engine structure  34  (or one or more of its components) or a preform thereof within the build space  140  in a layer-by-layer fashion. For example, the additive manufacturing apparatus  142  may deposit a first layer  144 A of powder over a support surface  146  within the build space  140 . The additive manufacturing apparatus  142  may thereafter selectively solidify (e.g., sinter) a select portion of the first layer  144 A of powder to form a first portion  148 A (e.g., layer, slice) of the stationary engine structure  34  (or one or more of its components) or a preform thereof. The additive manufacturing apparatus  142  may deposit a second layer  144 B of powder within the build space  140  over the first layer  144 A of at least partially solidified powder. The additive manufacturing apparatus  142  may thereafter selectively solidify a select portion of the second layer  144 B of powder with the previously solidified first portion  148 A to form a second portion  148 B (e.g., layer, slice) of the stationary engine structure  34  (or one or more of its components) or a preform thereof. This process may be repeated until the entire stationary engine structure  34  (or one or more of its components) or a preform thereof is formed. 
     While the additive manufacturing apparatus  142  is described above as a laser powder bed fusion (LPBF) apparatus, the present disclosure is not limited thereto. The additive manufacturing apparatus  142 , for example, may alternatively be configured as a stereolithography (SLA) apparatus, a direct selective laser sintering (DSLS) apparatus, an electron beam sintering (EBS) apparatus, an electron beam melting (EBM) apparatus, a laser engineered net shaping (LENS) apparatus, a laser net shape manufacturing (LNSM) apparatus, a direct metal deposition (DMD) apparatus or a direct metal laser sintering (DMLS) apparatus. 
       FIG.  7    is a flow diagram of a method  700  for manufacturing an engine such as, but not limited to, the gas turbine engine  20  (or one or more of its components) of  FIG.  1   . This method  700  may be performed using a manufacturing system such as, but not limited to, the manufacturing system  136  of  FIG.  6   . 
     In step  702 , a stationary engine structure preform  150  is formed. An example of the stationary engine structure preform  150  is schematically shown in  FIGS.  8 A and  8 B . This stationary engine structure preform  150  includes a plurality of engine structure section preforms such as, for example, an upstream section preform  152 , an inner downstream section preform  153  and an outer downstream section  154  preform. 
     Herein, the term preform may describe a body having at least a basic configuration (e.g., shape, size, features, etc.) of a part to be formed. For example, the upstream section preform  152  may generally have the same configuration as the upstream section  120  to be formed (see  FIGS.  1  and  3   ). The inner downstream section preform  153  may generally have the same configuration as the inner downstream section  121  to be formed (see  FIGS.  1  and  3   ). The outer downstream section preform  154  may generally have the same configuration as the outer downstream section  122  to be formed (see  FIGS.  3  and  5   ). However, one or more of these engine structure section preforms  152 - 154  may each include one or more build features, one or more unfinished surfaces, etc. The build features (e.g., webs, ribs, spares, etc.) may be included in order to support the respective preform during the formation step  702 . These build features, however, are removed (e.g., via machining) after the formation step  702  to provide the respective engine structure section  120 ,  121 ,  122 . In addition or alternatively, one or more surfaces, features (e.g., apertures, bosses, mounting features), etc. of the engine structure section preform may be machined (e.g., refined, smoothed, drilled, etc.) to provide the respective engine structure section  120 ,  121 ,  122 . Generally speaking, no additional material will need to be added to engine structure section preforms  120 - 122  to provide the respective engine structure sections  120 - 122  following the formation step  702 . 
     The stationary engine structure preform  150  may be completely formed within the build space  140  of the additive manufacturing apparatus  142 . The entire stationary engine structure preform  150  including its section preforms  152 - 154 , for example, may be additively manufactured concurrently using a layer-by-layer method within the build space  140 . Referring to  FIG.  8 A , the engine structure section preforms  152 - 154  may be formed as discrete (e.g., separate, distinct, unattached, etc.) bodies. Each engine structure section preform  152 ,  153 ,  154 , for example, may be separated from each adjacent engine structure section preform  152 ,  153 ,  154  by a gap. Alternatively, referring to  FIG.  8 B , one or more or all of the engine structure section preforms  152 - 154  may each be formed integral with and thereby connected to one or more or all of the other engine structure section preforms  152 - 154 . Each engine structure section preform  152 ,  153 ,  154  of  FIG.  8 B , for example, is connected to the adjacent engine structure section preform(s)  152 ,  153 ,  154  by an intermediate structure  156 A,  156 B (generally referred to as  156 ); e.g., a tie, a support, a bridge, etc. 
     While the section preforms  152 - 154  of the stationary engine structure preform  150  are described above as being formed concurrently within the build space  140 , the present disclosure is not limited thereto. For example, in other embodiments, each of the section preforms  152 - 154  (or select ones of the preforms) may be formed within the build space  140  separately from another one or more of the section preforms  152 - 154 . Furthermore, while the formation step  702  is described above in relation to additively manufacturing, the present disclosure is not limited thereto. For example, in other embodiments, one or more or each of the section preforms  152 - 154  may also or alternatively be formed using casting, machining and/or various other manufacturing techniques. 
     In step  704 , the stationary engine structure  34  and its engine structure sections  120 - 122  are provided. For example, the various preforms are removed from the build space  140  of the manufacturing system  136 . In addition (before or after being removed from the build space  140 ), one or more of the following operations may be performed:
         One or more or each of the engine structure section preforms  152 - 154  may be de-powdered. For example, any remaining, un-solidified powder trapped with a respective preform may be removed (e.g., evacuated) from that preform.   The engine structure section preforms  152 - 154  may be separated from one another where, for example, the preforms are connected together as shown in  FIG.  8 B . For example, the intermediate structures  156 A and/or  156 B of  FIG.  8 B  may be cut off and/or ground down to separate the engine structure section preforms.   Any build features may be removed from one or more or each of the engine structure section preforms  152 - 154 .   One or more surfaces, one or more features, etc. of one or more of the engine structure section preforms  152 - 154  may be machined, surface treated, heat treated, etc.   One or more surfaces of one or more of the engine structure section preforms  152 - 154  may be cleaned.   One or more surfaces of one or more of the engine structure section preforms  152 - 154  may be coated.
 
By performing one or more of the above operations and/or other (e.g., finishing) operations, the engine structure section preforms  152 - 154  may be turned into their respective engine structure sections  120 - 122 .
       

     In step  706 , the rotating assembly of the elements  42 ,  44  and  46  and the bearings  48  are installed with one or more of the engine structure sections  120  and/or  121 . 
     In step  708 , the engine structure sections  120 - 122  are mated and connected to provide the stationary engine structure  34  and, more generally, the gas turbine engine  20 . The upstream section  120 , for example, is axially abutted against and attached (e.g., mechanically fastened) to the inner downstream section  121 . The inner downstream section  121  is nested within the outer downstream section  122 , and the outer downstream section  122  is attached (e.g., bonded) to the inner downstream section  121 . 
     The gas turbine engine  20  is described above as a single spool, radial-flow turbojet turbine engine for ease of description. The present disclosure, however, is not limited to such an exemplary gas turbine engine. The gas turbine engine  20 , for example, may alternatively be configured as an axial flow gas turbine engine. The gas turbine engine  20  may be configured as a direct drive gas turbine engine. The gas turbine engine  20  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  20  may be configured with a single spool (e.g., see  FIG.  1   ), two spools, or with more than two spools. The gas turbine engine  20  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  20  is described above with an exemplary reverser flow annular combustor, the gas turbine engine  20  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 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.