Patent Publication Number: US-11649964-B2

Title: Fuel injector assembly for a turbine engine

Description:
BACKGROUND OF THE DISCLOSURE 
     1. Technical Field 
     This disclosure relates generally to a turbine engine and, more particularly, to a fuel injector assembly for the turbine engine. 
     2. Background Information 
     A combustor section in a modern a turbine engine includes one or more fuel injectors. Each fuel injector is operable to inject fuel for combustion within a combustion chamber. Various types and configurations of fuel injectors are known in the art. While these known fuel injectors have various benefits, there is still room in the art for improvement. There is a need in the art, for example, for fuel injectors with reduced manufacturing costs, that facilitate reduced assembly time as well as that reduce likelihood of carbon buildup within the combustion chamber caused by solidification of and/or traces of non-combusted fuel. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the present disclosure, an apparatus is provided for a turbine engine. This turbine engine apparatus includes a monolithic body. The monolithic body includes a splash plate and a fuel nozzle. The splash plate includes a splash plate surface. The fuel nozzle includes a nozzle orifice. The fuel nozzle is configured to direct fuel out of the nozzle orifice to impinge against the splash plate surface. 
     According to another aspect of the present disclosure, another apparatus is provided for a turbine engine. This turbine engine apparatus includes a structure, a fuel nozzle and a splash plate. The structure includes a fluid passage. The structure is configured to direct an axial fluid flow through the fluid passage. The fuel nozzle includes a nozzle orifice. The splash plate is arranged within the fluid passage and includes a splash plate surface. The fuel nozzle is configured to direct fuel out of the nozzle orifice to impinge against the splash plate surface. The splash plate is configured to disperse the fuel that impinges against the splash plate surface into the axial fluid flow. 
     According to still another aspect of the present disclosure, another apparatus is provided for a turbine engine. This turbine engine apparatus includes a fuel nozzle and a splash plate. The fuel nozzle includes a nozzle orifice. The splash plate includes a splash plate surface spaced from the fuel nozzle. The fuel nozzle is configured to direct a fuel jet out of the nozzle orifice along a fuel jet trajectory to the splash plate surface. The splash plate is configured to disperse fuel from the fuel jet in a radial outward pattern. The splash plate surface is angularly offset from the fuel jet trajectory by an acute angle. 
     The axial fluid flow may be or otherwise include a non-swirled fluid flow. 
     The splash plate may be integral with the fuel nozzle. 
     The splash plate may be configured with the fuel nozzle in a monolithic body. 
     The turbine engine assembly may also include a structure that includes an air passage. The structure may be configured to direct air through the air passage. The splash plate may be configured to disperse the fuel from the fuel jet in the radial outward pattern into the air within the air passage. 
     The fuel nozzle may be configured to direct the fuel out of the nozzle orifice as a fuel jet. The splash plate may be configured to redirect the fuel jet into a radiant pattern of fuel. 
     The splash plate may be spaced from and/or may overlap the nozzle orifice. 
     The splash plate surface may be configured as or otherwise include a planar splash plate surface. 
     The fuel nozzle may be configured to direct the fuel out of the nozzle orifice along a trajectory to impinge against the splash plate surface. The splash plate surface may be angularly offset from the trajectory by an acute angle. 
     The acute angle may be between sixty degrees (60°) and eighty degrees (80°). 
     The acute angle may be between thirty-five degrees (35°) and fifty-five degrees (55°). 
     The turbine engine assembly may also include a support member connecting and extending between the splash plate and the fuel nozzle. 
     The fuel nozzle may project into a flow passage. The support member may be upstream of the nozzle orifice relative to a fluid flow within the flow passage. 
     The turbine engine assembly may also include a second support member connecting and extending between the splash plate and the fuel nozzle. 
     The fuel nozzle may include a nozzle tube that has and extends along a longitudinal centerline. The nozzle orifice may be coaxial with the longitudinal centerline. 
     The turbine engine assembly may also include a fuel vaporizer. The splash plate may be configured to direct at least some of the dispersed fuel against the fuel vaporizer. 
     The turbine engine assembly may also include an air tube that includes an air passage. The fuel nozzle may project into the air passage. The splash plate may be arranged within the air passage such that the splash plate is configured to direct at least some of the dispersed fuel against an inner sidewall surface of the air tube. 
     The turbine engine assembly may also include a combustor wall at least partially forming a combustion chamber. The air tube may be connected to the combustor wall and/or may project into the combustion chamber. 
     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 
         FIGS.  1 - 4    are side sectional illustrations of portions of a turbine engine apparatus. 
         FIG.  5    is a perspective cross-sectional illustration of another portion of the turbine engine apparatus. 
         FIG.  6    is a side sectional illustration of another portion of the turbine engine apparatus. 
         FIG.  7    is a side cutaway illustration of another portion of the turbine engine apparatus schematically depicting an air flow and a fuel flow. 
         FIG.  8    is an illustration of a splash plate and a section of an associated support member further schematically depicting the fuel flow. 
         FIGS.  9  and  10    are perspective illustrations of the turbine engine apparatus configured with an additional support member for each splash plate. 
         FIG.  11    is a perspective cross-sectional illustration of a portion of a combustor section. 
         FIG.  12    is a perspective side sectional illustration of another portion of the combustor section. 
         FIG.  13    is a schematic side illustration of a single spool, radial-flow turbojet turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    illustrates a portion of an apparatus  20  for a turbine engine. This turbine engine apparatus  20  is configured as, or otherwise includes, a fuel injector assembly  22  for a combustor section of the turbine engine. The turbine engine apparatus  20  includes a fuel conduit  24 , a fuel nozzle  25  (e.g., a single and/or central orifice fuel nozzle) and a fuel nozzle splash plate  26 . The turbine engine apparatus  20  of  FIG.  1    may also include an apparatus base  27 , which apparatus base  27  may provide a structural support for the fuel conduit  24  and/or the fuel nozzle  25 . 
     The apparatus base  27  may be configured as any part of the turbine engine within the combustor section that is proximate the fuel injector assembly  22 . The apparatus base  27  of  FIG.  1   , for example, may be configured as a turbine engine case such as, but not limited to, a combustor section case, a diffuser case and/or a combustor wall. 
     The fuel conduit  24  is configured as, or may be part of, a fuel supply for the fuel nozzle  25 . The fuel conduit  24 , for example, may be or may be part of a fuel supply tube, a fuel inlet manifold and/or a fuel distribution manifold. The fuel conduit  24  is arranged at and/or is connected to a first side  30  (e.g., an exterior and/or outer side) of the apparatus base  27 . The fuel conduit  24  is configured with an internal fuel supply passage  32  formed by an internal aperture (e.g., a bore, channel, etc.) within the fuel conduit  24 . The supply passage  32  and the associated aperture extend within and/or through the fuel conduit  24  along a (e.g., curved or straight) centerline  34  of the supply passage  32 , which may also be a centerline of the fuel conduit  24 . 
     Referring to  FIG.  2   , the fuel nozzle  25  is configured to receive fuel from the fuel conduit  24 , and inject the received fuel into a plenum (e.g., a fluid passage  34  such as an air passage) at a distal end  36  (e.g., tip) of the fuel nozzle  25  to impinge against the splash plate  26 . The fuel nozzle  25  of  FIG.  2    includes a nozzle body  38  and a nozzle passage  40 ; e.g., a fuel passage. 
     The nozzle body  38  is arranged at and/or is connected to a second side  42  (e.g., an interior and/or inner side) of the apparatus base  27 , where the base second side  42  is opposite the base first side  30 . The nozzle body  38  of  FIG.  2    includes a nozzle tube  44  and a nozzle support structure  46  (e.g., a web). A base end of the nozzle tube  44  is connected to the apparatus base  27 . The nozzle tube  44  projects longitudinally out from the apparatus base  27  along a (e.g., straight or curved) longitudinal centerline  48  of the nozzle passage  40  and/or the nozzle tube  44  to the fuel nozzle distal end  36 . The nozzle support structure  46  is connected to and extends between the apparatus base  27  and a (e.g., upstream) side of the nozzle tube  44 . The nozzle support structure  46  structurally ties the nozzle tube  44  to the apparatus base  27  and may thereby support the nozzle tube  44  within the fluid passage  34 . The nozzle support structure  46 , for example, may form a support gusset for the nozzle tube  44 . 
     An internal bore of the nozzle tube  44  at least partially (or completely) forms the nozzle passage  40 . The nozzle passage  40  extends longitudinally along the longitudinal centerline  48  within and/or through the apparatus base  27  and the nozzle tube  44  from the supply passage  32  to a downstream nozzle orifice  50  at the fuel nozzle distal end  36 . This nozzle orifice  50  provides an outlet from the nozzle passage  40  and, more generally, the fuel nozzle  25 . 
     Referring to  FIG.  3   , the nozzle passage  40  includes one or more different flow portions (e.g.,  52 - 54 ) arranged longitudinally along the longitudinal centerline  48 . The nozzle passage  40  of  FIG.  3   , for example, includes a (e.g., upstream) flow channel portion  52 , a (e.g., intermediate) convergent portion  53  and a (e.g., downstream) throat portion  54 . 
     The flow channel portion  52  is upstream of the convergent portion  53 , for example at (e.g., on, adjacent or proximate) an upstream end of the nozzle passage  40 . The flow channel portion  52  of  FIG.  3   , for example, is formed by a (e.g., non-tapering, cylindrical) flow channel sidewall surface  56 . This flow channel sidewall surface  56  and, thus, the flow channel portion  52  extends longitudinally along the longitudinal centerline  48  from the supply passage  32  to the convergent portion  53 . 
     The convergent portion  53  is fluidly coupled between the flow channel portion  52  and the throat portion  54 . The convergent portion  53  of  FIG.  3   , for example, is formed by a tapering (e.g., frustoconical) convergent sidewall surface  58 . This convergent sidewall surface  58  and, thus, the convergent portion  53  extends longitudinally along the longitudinal centerline  48  from the flow channel portion  52  to the throat portion  54 , where a width  60  (e.g., diameter) of the convergent sidewall surface  58  decreases as the convergent portion  53  extends longitudinally towards the throat portion  54 /the nozzle orifice  50 . 
     The throat portion  54  is downstream of the convergent portion  53  and/or at the nozzle orifice  50 , for example at (e.g., on, adjacent or proximate) the fuel nozzle distal end  36 . The throat portion  54  of  FIG.  3   , for example, is formed by a (e.g., non-tapering, cylindrical) throat sidewall surface  62 . This throat sidewall surface  62  and, thus, the throat portion  54  extends longitudinally along the longitudinal centerline  48  from the convergent portion  53  to (or towards) the nozzle orifice  50 . A downstream most end of the throat portion  54  may thereby define the nozzle orifice  50 . Of course, in other embodiments, the nozzle passage  40  may also include another flow portion (e.g., a divergent portion) arranged longitudinally between the throat portion  54  and the nozzle orifice  50 . In still other embodiments, any one or more of the foregoing flow portions  52 - 54  may also or alternatively be omitted; e.g., the flow channel portion  52  may be omitted where, for example, the convergent portion  53  extends from the supply passage  32  to the throat portion  54 . The present disclosure therefore is not limited to the foregoing exemplary nozzle passage configurations. 
     Referring to  FIG.  4   , the splash plate  26  is configured to redirect (e.g., disperse) the fuel injected into the fluid passage  34  from the fuel nozzle  25  into a disperse (e.g., a widespread) pattern (e.g., see  FIGS.  7  and  8   ). The splash plate  26 , for example, is arranged proximate and laterally overlaps the nozzle orifice  50 . The splash plate  26  is longitudinally spaced from the fuel nozzle  25  and its nozzle orifice  50  by a longitudinal distance  64  along the longitudinal centerline  48 . This longitudinal distance  64  may be equal to or different (e.g., greater or less) than a width (e.g., diameter) of the nozzle passage  40 . The longitudinal distance  64  of  FIG.  4   , for example, is between one-half times (0.5×) and five times (5×) a width  66  (e.g., a diameter) of the throat portion  54 . The present disclosure, however, is not limited to the foregoing exemplary dimensional relationship between the splash plate  26  and the fuel nozzle  25 . 
     The splash plate  26  of  FIGS.  4  and  5    is configured with a (e.g., circular) puck-like body. The splash plate  26  of  FIG.  4   , for example, extends axially along a centerline axis  68  of the splash plate  26  between a frontside splash plate surface  70  and a backside splash plate surface  72 , which backside splash plate surface  72  is axially opposite the frontside splash plate surface  70 . Each of these splash plate surfaces  70  and  72  may have a generally circular shape. However, in other embodiments, one or more of the splash plate surfaces  70  and  72  may each have a non-circular (e.g., oval, polygonal, etc.) shape. Each of the splash plate surfaces  70  and  72  may be configured as a smooth and/or planar surface. However, in other embodiments, one or more of the splash plate surfaces  70  and  72  may each be configured as a non-planar (e.g., concave, convex, etc.) surface and/or with one or more flow disruptions; e.g., apertures or projections. The splash plate  26  of  FIGS.  4  and  5    also includes at least one side perimeter surface  74  that extends axially between the opposing splash plate surfaces  70  and  72  and circumferentially about the centerline axis  68  of the splash plate  26 . 
     Referring to  FIG.  4   , the splash plate  26  and, more particularly, its frontside splash plate surface  70  is angularly offset from the longitudinal centerline  48  and/or fuel trajectory  90  (discussed below) by a first acute angle  76  (an angle that is greater than zero degrees and less than ninety degrees) when viewed, for example, in the plane of  FIG.  4   ; e.g., a plane that laterally bisects one or more or each of the components  26 ,  44  and  46  and/or is parallel with and coincident with the centerline  48 . The first acute angle  76  may be between sixty degrees (60°) and eighty degrees (80°) as shown in  FIG.  4   ; e.g., the first acute angle  76  may be substantially (e.g., +/−2°) or exactly equal to seventy degrees (70°). In another example, the first acute angle  76  may be between thirty-five degrees (35°) and fifty-five degrees (55°) as shown in  FIG.  6   ; e.g., the first acute angle  76  may be substantially (e.g., +/−2°) or exactly equal to forty-five degrees (45°). 
     The splash plate  26  of  FIG.  4    and, more particularly, its frontside splash plate surface  70  is angularly offset from a plane of the nozzle orifice  50  and/or a surface  78  of the nozzle tube  44  at the fuel nozzle distal end  36  by a second acute angle  80 . The second acute angle  80  may be between ten degrees (10°) and thirty degrees (30°) as shown in  FIG.  4   ; e.g., the second acute angle  80  may be substantially (e.g., +/−2°) or exactly equal to twenty degrees (20°). In another example, the second acute angle  80  may be between thirty-five degrees (35°) and fifty-five degrees (55°) as shown in  FIG.  6   ; e.g., the second acute angle  80  may be substantially (e.g., +/−2°) or exactly equal to forty-five degrees (45°). 
     The splash plate  26  of  FIGS.  4  and  5    is connected to the fuel nozzle  25  by at least (or only) one support member  82 . The support member  82  may be configured as a beam and/or a pylon. The support member  82  of  FIGS.  4  and  5   , for example, has an elongated body that is connected to and extends between the fuel nozzle  25  and the splash plate  26 . More particularly, the support member  82  of  FIGS.  4  and  5    is connected (e.g., directly) to and extends between the nozzle support structure  46  and the splash plate  26 . Of course, in other embodiments, the support member  82  may also or alternatively be connected to and/or project out from the nozzle tube  44 . 
     The support member  82  of  FIG.  4    is arranged at (e.g., on, adjacent or proximate) an upstream end  84  of the splash plate  26  relative to a fluid flow  85  (e.g., an air flow) within the fluid passage  34  (e.g., an air passage). The support member  82  may thereby be arranged upstream of the nozzle orifice  50  relative to the fluid flow  85  within the fluid passage  34 . With such an arrangement, the fuel redirected (e.g., dispersed) by the splash plate  26  may flow unobstructed in a downstream direction through a spatial gap  86  between the splash plate  26  and the fuel nozzle  25 . The present disclosure, however, is not limited to such an exemplary support member placement. 
     Referring to  FIG.  2   , during turbine engine operation, fuel is directed into the supply passage  32  from a fuel source (not shown). At least a portion (or all) of the fuel within the supply passage  32  is directed into the nozzle passage  40 . Referring to  FIG.  7   , this fuel flows through the nozzle passage  40  and out of the fuel nozzle  25  through the nozzle orifice  50  and into the fluid passage  34  (more particularly, into the spatial gap  86 ) as a fuel jet  88  along a fuel jet trajectory  90 , which may be parallel (e.g., coaxial) with the centerline  48 . This fuel jet  88  may be a linear concentrated flow/stream of fuel versus, for example, a spread-out pattern of fuel such as a conical film of fuel. The fuel jet  88  flows through the spatial gap  86  along its trajectory  90  and impacts (e.g., impinges against) the frontside splash plate surface  70  at a target area; e.g., an impingement area. Referring to  FIGS.  7  and  8   , upon impacting the frontside splash plate surface  70 , the splash plate  26  redirects (e.g., disperses) the impinging fuel jet  88  radially outward (relative to the fuel jet trajectory  90 ) into a (e.g., uniform and/or symmetrical) disperse radiant pattern  92  (e.g., an arcuate and/or a generally planar film; schematically shown in  FIGS.  7  and  8    via discrete flow arrows). The fuel may thereby be more evenly dispersed/spread/mixed into the fluid (e.g., air) flowing past the fuel nozzle  25  and the splash plate  26  within the fluid passage  34 . Providing such relatively even mixing of the fuel and the fluid may in turn increase fuel burn efficiency and/or reduce likelihood of carbon formation within the turbine engine. 
     In some embodiments, referring to  FIG.  2   , the splash plate  26  is cantilevered from the fuel nozzle  25  through the support member  82 . In other embodiments, the splash plate  26  may be further supported by at least one additional support member  82 B as shown, for example, in  FIGS.  9  and  10   . This downstream support member  82 B is connected to and extends between the fuel nozzle  25  and the splash plate  26 . More particularly, the downstream support member  82 B of  FIG.  9    is connected to and projects out from the nozzle tube  44 . The downstream support member  82 B of  FIG.  10    is connected to and projects out from another (e.g., downstream) nozzle support structure  46 B (e.g., web) for the fuel nozzle  25 . Referring to  FIGS.  9  and  10   , the downstream support member  82 B may be positioned opposite to (e.g., diametrically opposed with) the upstream support member  82 ; however, the present disclosure is not limited to such exemplary support member locations. 
     In addition to increasing structural ties between the splash plate  26  and the fuel nozzle  25 , including more than one support member (e.g.,  82 ,  82 B) may also provide for reducing the size of the support member (e.g.,  82 ,  82 B) e.g., thickness. Reducing the size of the support member(s) (e.g.,  82 ,  82 B) may in turn reduce flow impedance to the dispersed fuel traveling past the support members (e.g.,  82 ,  82 B) and, thus, promote further mixing between the fuel and the fluid flow; e.g., air flow. 
     In some embodiments, referring to  FIG.  11   , the fuel nozzle  25  may be one of a plurality of fuel nozzles  25  connected to the apparatus base  27  and fluidly coupled with the fuel conduit  24 . These fuel nozzles  25  may be arranged circumferentially about a centerline/rotational axis  94  of the turbine engine in an annular array. Each of the fuel nozzles  25  may be associated with a respective splash plate  26 . 
     In some embodiments, referring to  FIG.  2   , the apparatus base  27 , the fuel conduit  24  and each fuel nozzle  25  may be configured together in an integral, monolithic body. Each fuel nozzle  25  and its respective splash plate  26  may also or alternatively be configured together in the monolithic body. In such embodiments, selecting the first acute angle  76  of  FIG.  4    to be between sixty degrees and eighty degrees (e.g., approximately or exactly seventy degrees) and/or the second acute angle  80  to be between ten degrees and thirty degrees (e.g., approximately or exactly twenty degrees) may facilitate additive manufacturing of the turbine engine apparatus  20  as a monolithic body. The present disclosure, however, is not limited to such an exemplary construction. For example, in other embodiments, one or more or each of the apparatus components and/or portions thereof may be individually formed and subsequently connected (e.g., fastener and/or bonded) together. 
     In some embodiments, referring to  FIGS.  11  and  12   , the turbine engine apparatus  20  may also include one or more fuel vaporizers  96 . Each fuel nozzle  25  is arranged with a respective one of the fuel vaporizers  96 . More particularly, each fuel nozzle  25  projects into a respective one of the fuel vaporizers  96  and the associated splash plate  26  is arranged within a fluid passage (e.g., an air passage) of the respective fuel vaporizer  96 . With such an arrangement, each splash plate  26  may direct a portion of the dispersed fuel to impinge against a surface  98  of the respective fuel vaporizer  96 . The fuel vaporizer  96  may provide initial or further vaporization of the dispersed fuel. Each splash plate  26  may also direct another portion of the dispersed fuel to mix with the passing fluid (e.g., air) without impinging against the fuel vaporizer  96 . 
     The ratio of an amount of the dispersed fuel which contacts the fuel vaporizer  96  versus an amount of the dispersed fuel which does not contact the fuel vaporizer  96  may be controlled by adjusting a value of the first acute angle  76  of  FIGS.  4  and  6   . For example, when the value of the first acute angle  76  is increased towards ninety degrees (e.g., see  FIG.  4   ), more of the fuel dispersed by the splash plate  26  may penetrate further into the fluid flow and, thus, more of the dispersed fuel may contact the fuel vaporizer  96  (see  FIG.  12   ). By contrast, when the value of the first acute angle  76  is decreased towards zero degrees (e.g., see  FIG.  6   ), less of the fuel dispersed by the splash plate  26  may penetrate far into the fluid flow and, thus, less of the dispersed fuel may contact the fuel vaporizer  96  (see  FIG.  12   ). 
     In the specific embodiment of  FIGS.  11  and  12   , each fuel vaporizer  96  is configured as a structure such as a flow tube  100  (e.g., a fluid tube, an air tube) for a combustor  102  in the combustor section  104 . Note, the combustor  102  may also include at least one flow tube  106  in between, for example, each circumferentially neighboring set of the vaporizers  96  and/or one or more flow tubes  108  in another (e.g., forward/upstream) array. Each of the flow tubes  100 ,  106 ,  108  is connected to and projects out from a wall  110  of the combustor  102  and into a combustion chamber  112  at least partially defined by the combustor wall  110 . The fluid passage  34  (e.g., air passage) of each flow tube  100  is configured to receive fluid and, more particularly, compressed air from a compressor section of the turbine engine (not visible in  FIGS.  11  and  12   ) through another plenum  114 . This compressed air is directed through the respective fluid passage  34  and into the combustion chamber  112 . However, before reaching the combustion chamber  112 , the air within the respective fluid passage  34  is mixed with fuel dispersed from a respective one of the splash plates  26  to provide a mixture of compressed air and atomized fuel. By dispersing the fuel within the flow tube  100 , the fuel may be more likely to atomize within the respective fluid passage  34 ; e.g., upon dispersing into the airflow and/or upon impinging against the surface  98  (e.g., an inner side wall surface of the flow tube  100 ). By increasing atomization of the fuel, the fuel injector assembly  22  may reduce the likelihood of carbon buildup within the fluid passage  34  and/or within the combustion chamber  112 . 
     In some embodiments, each fuel vaporizer  96 /flow tube  100  is configured to direct an axial fluid flow therewith/therethrough. The term axial fluid flow may describe a straight or linear flow of fluid such as a non-swirled fluid flow; e.g., non-swirled air. For example, none of the fuel vaporizers  96 /flow tubes  100  is configured with or otherwise receives its fluid (e.g., air) directly and/or indirectly from a swirler. Thus, the fluid flowing through each fuel vaporizer  96 /flow tube  100  is non-swirled; e.g., the fluid primarily (or only) has axial velocity/momentum components with little or no tangential velocity/momentum components. Of course, the fluid flowing through each fuel vaporizer  96 /flow tube  100  may include relatively low level flow disruptions, turbulence, vortices, etc. caused when, for example, the fluid turns from the plenum  114  into the fluid passage  34 , etc. 
     The turbine engine apparatus  20  of the present disclosure may be configured with various different types and configurations of turbine engines.  FIG.  13    illustrates one such type and configuration of the turbine engine—a single spool, radial-flow turbojet turbine engine  116  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.  13   , the turbine engine  116  includes an upstream inlet  118 , a (e.g., radial) compressor section  120 , the combustor section  104 , a (e.g., radial) turbine section  122  and a downstream exhaust  124  fluidly coupled in series. A compressor rotor  126  in the compressor section  120  is coupled with a turbine rotor  128  in the turbine section  122  by a shaft  130 , which rotates about the centerline/rotational axis  94  of the turbine engine  116 . 
     The turbine engine apparatus  20  may be included in various turbine engines other than the one described above. The turbine engine apparatus  20 , for example, may be included in a geared turbine engine where a gear train connects one or more shafts to one or more rotors in a fan section, a compressor section and/or any other engine section. Alternatively, the turbine engine apparatus  20  may be included in a turbine engine configured without a gear train. The turbine engine apparatus  20  may be included in a geared or non-geared turbine engine configured with a single spool (e.g., see  FIG.  13   ), with two spools, or with more than two spools. The turbine engine may be configured as a turbofan engine, a turbojet engine, a propfan engine, a pusher fan engine or any other type of turbine engine. The present disclosure therefore is not limited to any particular types or configurations of turbine engines. The present disclosure is also not limited to a propulsion system application. For example, the gas turbine engine may alternatively be configured as an auxiliary power unit (APU) or an industrial gas turbine engine. 
     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.