Patent Publication Number: US-11021961-B2

Title: Rotor assembly thermal attenuation structure and system

Description:
FIELD 
     The present subject matter relates generally to rotor assembly thermal attenuation and flow structures for heat engines. 
     BACKGROUND 
     Heat engines, such as gas turbine engines, generally include cooling structures to provide cooling fluid to turbine blades to reduce wear and deterioration. However, known structures and systems for providing cooling fluid to turbine blades often result in inefficiencies due to large pressure drops and high temperatures related to the cooling fluid and the cooling fluid source. As such, there is a need for structures and systems for improving provision of cooling fluid to turbine blades while mitigating losses and inefficiencies at the engine related to providing cooling fluid. 
     BRIEF DESCRIPTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     An aspect of the present disclosure is directed to a rotor assembly for a turbine engine. The rotor assembly includes an airfoil assembly and a hub to which the airfoil assembly is attached. A wall assembly defines a first cavity and a second cavity between the airfoil assembly and the hub. The first cavity and the second cavity are at least partially fluidly separated by the wall assembly. The first cavity is in fluid communication with a flow of first cooling fluid and the second cavity is in fluid communication with a flow of second cooling fluid different from the first cooling fluid. 
     In one embodiment, the wall assembly is extended from the airfoil assembly or the hub to define a seal assembly defining the first cavity and the second cavity. 
     In another embodiment, the wall assembly is extended from the airfoil assembly between a static assembly and the rotor assembly to define a plenum therewithin in fluid communication with one or more of the first cavity or the second cavity. 
     In various embodiments, the rotor assembly includes a wall within the airfoil assembly defining a first plenum fluidly separated from a second plenum. In one embodiment, the first plenum is in fluid communication with the first cavity, and the second plenum is in fluid communication with the second cavity. 
     In one embodiment, the rotor assembly defines a first inlet opening through a base portion of the airfoil assembly in fluid communication with the first cavity. 
     In various embodiments, the airfoil assembly includes a plurality of circuits in fluid communication with one or more of the first cavity and the second cavity. In one embodiment, the plurality of circuits includes a first circuit in fluid communication with the first cavity and a third circuit in fluid communication with the second cavity. In another embodiment, the plurality of circuits includes a second circuit in fluid communication with the first cavity. In yet another embodiment, the plurality of circuits includes a second circuit in fluid communication with the second cavity. 
     Another aspect of the present disclosure is directed to a heat engine. The heat engine includes a first cooling fluid source configured to provide a first cooling fluid; a second cooling fluid source configured to provide a second cooling fluid, wherein the first cooling fluid and the second cooling fluid each define one or more of a different pressure or temperature relative to one another; and a rotor assembly including an airfoil assembly and a hub to which the airfoil assembly is attached. The rotor assembly defines a first cavity and a second cavity between the airfoil assembly and the hub at least partially fluidly separates the first cavity from the second cavity. The first cavity is in fluid communication with the first cooling fluid source to receive the first cooling fluid. The second cavity is in fluid communication with the second cooling fluid source to receive the second cooling fluid. 
     In various embodiments, the heat engine further includes a first static assembly disposed directly adjacent to the rotor assembly. The first cooling fluid source is disposed at least partially through the first static assembly. The first cooling fluid source is configured to provide the first cooling fluid therethrough to the first cavity of the rotor assembly. The heat engine further includes a second static assembly disposed directly adjacent to the rotor assembly. The second cooling fluid source is disposed at least partially through the second static assembly. The second cooling fluid source is configured to provide the second cooling fluid therethrough to the second cavity of the rotor assembly. 
     In one embodiment, the rotor assembly includes a wall defining a first plenum fluidly separated from a second plenum. The first plenum is in fluid communication with the first cavity. The second plenum is in fluid communication with the second cavity. 
     In another embodiment, the wall assembly is extended from a base portion of the airfoil assembly and the hub to define a seal assembly defining the first cavity and the second cavity between the airfoil assembly and the hub. 
     In yet another embodiment, the wall assembly is extended from the airfoil assembly between the rotor assembly and one or more of the first static assembly or the second static assembly to define one or more of the first plenum or the second plenum therewithin. 
     In one embodiment, the rotor assembly defines a first inlet opening through the base portion in fluid communication with the first cavity. 
     In various embodiments, the rotor assembly includes a plurality of circuits through the airfoil assembly in fluid communication with one or more of the first cavity and the second cavity. In one embodiment, the plurality of circuits through the rotor assembly includes a first circuit in fluid communication with the first cavity and a third circuit in fluid communication with the second cavity. In another embodiment, the plurality of circuits through the rotor assembly includes a second circuit in fluid communication with the first cavity. In yet another embodiment, the plurality of circuits includes a second circuit in fluid communication with the second cavity. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  is a schematic cross sectional view of an exemplary heat engine including a rotor assembly according to aspects of the present disclosure; 
         FIG. 2  is a schematic cross sectional view of an exemplary portion of a turbine section and combustion section of the engine of  FIG. 1 ; 
         FIG. 3  is a detailed schematic cross sectional view of an exemplary embodiment of a portion of the turbine section and combustion section of  FIG. 2 ; 
         FIG. 4  is a detailed schematic cross sectional view of another exemplary embodiment of a portion of the turbine section and combustion section of  FIG. 2 ; 
         FIG. 5  is a perspective view of an exemplary embodiment of an airfoil assembly of the rotor assembly depicted in regard to  FIGS. 1-4 ; 
         FIG. 6  is a cross sectional view of an exemplary embodiment of the airfoil assembly of  FIG. 5 ; 
         FIG. 7  is another cross sectional view of an exemplary embodiment of the airfoil assembly of  FIG. 5 ; 
         FIG. 8  is a schematic cross sectional view of an exemplary embodiment of the airfoil assembly of  FIGS. 5-7 ; 
         FIG. 9  is a schematic cross sectional view of another exemplary embodiment of the airfoil assembly of  FIGS. 5-7 ; 
         FIG. 10  is a schematic cross sectional view of yet another exemplary embodiment of the airfoil assembly of  FIGS. 5-7 ; 
         FIG. 11  is a schematic cross sectional view of still another exemplary embodiment of the airfoil assembly of  FIGS. 5-7 ; and 
         FIG. 12  is a schematic cross sectional view of still yet another exemplary embodiment of the airfoil assembly of  FIGS. 5-7 ; 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     Approximations recited herein may include margins based on one more measurement devices as used in the art, such as, but not limited to, a percentage of a full scale measurement range of a measurement device or sensor. Alternatively, approximations recited herein may include margins of 10% of an upper limit value greater than the upper limit value or 10% of a lower limit value less than the lower limit value. 
     Embodiments of an engine including a rotor assembly and airfoil assembly are generally provided that may improve provision of cooling fluid to rotor blades while mitigating losses and inefficiencies at the engine related to providing cooling fluid. Embodiments shown and described herein include providing two or more cooling fluids of different pressure and/or temperatures to forward and aft portions of the rotor assembly. The different cooling fluids may generally include a cooled cooling air (CCA) circuit such as to pass compressor section air through one or more heat exchangers and through a static structure such as to provide cooling fluid to the airfoil assembly of the rotor assembly. The other fluid may generally include a higher pressure and/or higher temperature source, such as routed through the combustion section. The separate flows of cooling fluid reduce the overall flow of cooling fluid extracted from the aerodynamic and thermodynamic cycle of the engine via reducing the flow extracted through the combustion section and providing a reduced flow of lower temperature cooling fluid through the rotor assembly. 
     Referring now to the drawings,  FIG. 1  is a schematic partially cross-sectioned side view of an exemplary heat engine  10  herein referred to as “engine  10 ” as may incorporate various embodiments of the present disclosure. Although further described below with reference to a gas turbine engine, the present disclosure is also applicable to turbomachinery in general, including gas turbine engines defining turbofan, turbojet, turboprop, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units, and steam turbine engines, internal combustion engines, reciprocating engines, and Brayton cycle machines generally. As shown in  FIG. 1 , the engine  10  has a longitudinal or axial centerline axis  12  that extends there through for reference purposes. In general, the engine  10  may include a fan assembly  14  and a core engine  16  disposed downstream from the fan assembly  14 . 
     The core engine  16  may generally include a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  encases or at least partially forms, in serial flow relationship, a compressor section  21  having a booster or low pressure (LP) compressor  22 , a high pressure (HP) compressor  24 , a combustor-diffuser assembly  26 , a turbine section  31  including a high pressure (HP) turbine  28 , a low pressure (LP) turbine  30  and a jet exhaust nozzle section  32 . A high pressure (HP) rotor shaft  34  drivingly connects the HP turbine  28  to the HP compressor  24 . A low pressure (LP) rotor shaft  36  drivingly connects the LP turbine  30  to the LP compressor  22 . The LP rotor shaft  36  may also be connected to a fan shaft  38  of the fan assembly  14 . In particular embodiments, as shown in  FIG. 1 , the LP rotor shaft  36  may be connected to the fan shaft  38  by way of a reduction gear  40  such as in an indirect-drive or geared-drive configuration. In other embodiments, the engine  10  may further include an intermediate pressure (IP) compressor and turbine rotatable with an intermediate pressure shaft. 
     As shown in  FIG. 1 , the fan assembly  14  includes a plurality of fan blades  42  that are coupled to and that extend radially outwardly from the fan shaft  38 . An annular fan casing or nacelle  44  circumferentially surrounds the fan assembly  14  and/or at least a portion of the core engine  16 . In one embodiment, the nacelle  44  may be supported relative to the core engine  16  by a plurality of circumferentially-spaced outlet guide vanes or struts  46 . Moreover, at least a portion of the nacelle  44  may extend over an outer portion of the core engine  16  so as to define a bypass airflow passage  48  therebetween. 
     It should be appreciated that the HP turbine  28 , the HP shaft  34 , and the HP compressor  24  together may define a rotor assembly  90  of the engine  10  rotatable relative to the centerline axis  12 . In other embodiments, the rotor assembly  90  further described herein may include the LP turbine  30 , the LP shaft  36 , and the LP compressor  22  together, or, alternatively, including the fan shaft  38 . In still other embodiments not depicted, the rotor assembly  90  may include an intermediate pressure turbine, shaft, and compressor assembly. 
     During operation of the engine  10 , a volume of oxidizer as indicated schematically by arrows  74  enters the engine  10  through an associated inlet  76  of the nacelle  44  and/or fan assembly  14 . As the oxidizer  74  passes across the fan blades  42  a portion of the oxidizer as indicated schematically by arrows  78  is directed or routed into the bypass airflow passage  48  while another portion of the oxidizer as indicated schematically by arrow  80  is directed or routed into the LP compressor  22 . Oxidizer  80  is progressively compressed as it flows through the LP and HP compressors  22 ,  24  towards the combustion section  26 . 
     Combustion gases  86  generated at the combustion section  26  flow into the turbine section  31 , such as to the HP turbine  28 , thus causing the HP rotor shaft  34  to rotate, thereby supporting operation of the HP compressor  24 . As shown in  FIG. 1 , the combustion gases  86  are then routed through the LP turbine  30 , thus causing the LP rotor shaft  36  to rotate, thereby supporting operation of the LP compressor  22  and/or rotation of the fan shaft  38 . The combustion gases  86  are then exhausted through the jet exhaust nozzle section  32  of the core engine  16  to provide propulsive thrust. 
     Typically, the LP and HP compressors  22 ,  24  provide more oxidizer to the combustion section  26  than is utilized for producing combustion gases  86 . Therefore, a portion of the oxidizer  82  as indicated schematically by arrows  83  may be used as a first cooling fluid. For example, as shown in  FIG. 2 , the first cooling fluid  83  may be routed through a first conduit  66  to provide thermal attenuation (e.g., heat transfer generally, or cooling specifically) to hotter portions of the rotor assembly  90 , such as at the HP turbine  28  and/or LP turbine  30 . In various embodiments, the first conduit  66  is defined at the combustion section  26  and/or turbine section  31 , such as depicted in part at least at  FIG. 2 . The first conduit  66  may generally provide the first cooling fluid  83  via one or more walls  301  defining a passage  65  between the wall  301  and at least one component at the rotor assembly  90 . The first conduit  66  is in fluid communication with a first cavity  116  ( FIGS. 3-7 ) at the rotor assembly  90  such as to provide a flow of the first cooling fluid  83  to the rotor assembly  90  such as further described below in regard to  FIGS. 3-12 . 
     The engine  10  may generally include a first static assembly  310  disposed adjacent to the rotor assembly  90  along an axial direction A, such as directly forward of the rotor assembly  90 . The first static assembly  310  may include the combustion section  26  upstream of the HP turbine  28  including the rotor assembly  90 . Still further, the first static assembly  310  may define, at least in part, the first conduit  66  through which the first cooling fluid  83  from a first cooling fluid source  200  is provided to the first cavity  116  ( FIGS. 3-7 ) of the rotor assembly  90 . 
     Referring still to  FIG. 2 , the first cooling fluid  83  through the first conduit  66  may generally be provided by a first cooling fluid source  200  configured to provide the first cooling fluid  83 . In various embodiments, the first cooling fluid source  200  may define one or more portions of the compressor section  21 , such as form a compressor bleed at the LP compressor  22  or HP compressor  24 . In one embodiment, the first cooling fluid source  200  is defined at the exit of the compressor section  21  (e.g., at the combustion section  26 ). In various embodiments, the first cooling fluid source  200  is defined from one or more stages within the compressor section  21  upstream of a compressor exit  64  ( FIG. 1 ). 
     In various embodiments, the engine  10  further includes a second cooling fluid source  300  configured to provide a second cooling fluid from a portion of the flow of oxidizer  82 , such as depicted via arrows  84 . The second cooling fluid source  300  may additionally derive the second cooling fluid  84  from the compressor section  21 . However, the second cooling fluid source  300  may further include one or more flow paths defining the second cooling fluid  84  of one or more of a different pressure or temperature relative to the first cooling fluid  83 . In various embodiments, the second cooling fluid source  300  may further include one or more heat exchangers. For example, the second cooling fluid source  300  may provide the second cooling fluid  84  in thermal communication with one or more of a flow of bypass air (e.g., flow of oxidizer  78 ), a flow of liquid and/or gaseous fuel, a flow of lubricant, a flow of hydraulic fluid, a flow of cryogenic fluid, supercritical fluid, or other coolant or refrigerant, or other heat sink, such as to decrease the temperature of the second cooling fluid  84  relative to the flow of oxidizer  82 . 
     The engine  10  may generally include a second static assembly  320  disposed adjacent to the rotor assembly  90  along the axial direction A, such as directly aft of the rotor assembly  90 . The second static assembly  320  may include a portion of the HP turbine  28 , such as a casing, frame, or vane assembly, downstream of one or more rotors of the turbine section  31 . Still further, the second static assembly  320  may define, at least in part, a second passage  67  through which the second cooling fluid  84  from the second cooling fluid source  300  is provided to a second cavity  117  ( FIGS. 3-7 ) of the rotor assembly  90 , such as further described herein. 
     Referring now to  FIGS. 2-3 , schematic cross sectional views of the engine  10  are generally provided.  FIGS. 2-3  generally depict portions of the turbine section, such as the HP turbine  28 , and an exit portion of the combustion section  26 , such as at the turbine nozzle assembly  68 . The engine  10  includes the rotor assembly  90  including an airfoil assembly  100  and a hub  140  to which the airfoil assembly  100  is attached. The airfoil assembly  100  includes a base portion  110  coupled to the hub  140 . In various embodiments, the airfoil assembly  100  is detachably coupled to the hub  140 . For example, the hub  140  may define a slot, such as a dovetail slot through which the airfoil assembly  100  may be detachably coupled. However, in other embodiments, the airfoil assembly  100  may be integral to the hub  140 , such as defining an integrally bladed rotor or bladed disk. 
     Referring to  FIG. 3 , the rotor assembly  90  may include a seal assembly  130  extended from the base portion  110  of the airfoil assembly  100  to the hub  140 . The seal assembly  130  defines a first cavity  116  and a second cavity  117  separated from one another by the seal assembly  130 . In various embodiments, the first cavity  116  and the second cavity  117  are defined collectively by the hub  140 , the base portion  110 , and the seal assembly  130 . The seal assembly  130  fluidly separates the first cavity  116  and the second cavity  117  between the airfoil assembly  100  and the hub  140 . For example, the seal assembly  130  enables the fluidly separate flows of cooling fluids  83 ,  84  to enter into the base portion  110  of the airfoil assembly  100  from their respective cavities  116 ,  117 , such as further depicted in regard to  FIGS. 8-12 . In various embodiments, the seal assembly  130  may define a labyrinth seal, a brush seal, a leaf seal, a foil or other single or multi-walled seal, or other appropriate sealing arrangement. 
     In various embodiments, the seal assembly  130  includes a wall assembly  135  coupled to the rotor assembly  90 . The wall assembly  135  is coupled to airfoil assembly  100  and extended therefrom to fluidly separate the flows of cooling fluid  83 ,  84  from one another. Referring to  FIG. 3 , in one embodiment, the seal assembly  130  including the wall assembly  135  is coupled to the base portion  110  of the airfoil assembly  100 . The wall assembly  135  defines the first cavity  116  fluidly segregated from the second cavity  117 . It should be appreciated that the seal assembly  130  separates or disconnects fluid flow between the first cavity  116  and the second cavity  117 . However, in various embodiments, a quantity of flow may flow between the first cavity  116  and the second cavity  117 . 
     In various embodiments, such as depicted in regard to  FIGS. 3-4 , the wall assembly  135  includes a first wall  131  extended from the base portion  110  of the airfoil assembly  100  and in contact with the hub  140 . In another embodiment, such as depicted in regard to  FIG. 3 , the wall assembly  135  further includes a second wall  132  extended from the hub  140  in contact with the base portion  110  of the airfoil assembly  100 . The first wall  131  and the second wall  132  are in direct adjacent arrangement such as to provide a sealing arrangement fluidly disconnecting the first cavity  116  and the second cavity  117 . For example, the first wall  131  and the second wall  132  may each be in direct adjacent arrangement along a chordwise direction  91  ( FIG. 3 ) relative to the airfoil assembly  100 . The seal assembly  130  may further include an alternating plurality of the first wall  131  and the second wall  132  such as to define cavities therebetween to limit flow or fluid communication between the first cavity  116  and the second cavity  117 . 
     Referring back to  FIG. 3 , in various embodiments, the seal assembly  130  defines the first cavity  116  between the base portion  110  and the hub  140  along the radial direction R. In another embodiment, the seal assembly  130  defines the second cavity  117  between the base portion  110  and the hub  140  along the radial direction R. In still various embodiments, a first inlet opening  111  and a second inlet opening  112  are each separated by the seal assembly  130  therebetween. In various embodiments, the first inlet opening  111  and the second inlet opening  112  are separated by the seal assembly  130  along the chordwise direction  91  corresponding to the axial direction A of the engine  10 . In one embodiment, the base portion  110  defines the first inlet opening  111  in direct fluid communication with the first cavity  116 . In another embodiment, the second inlet opening  112  is defined in direct fluid communication with the second cavity  117 . 
     Referring now to  FIG. 4 , another exemplary embodiment of the engine  10  is generally provided. The embodiment provided in regard to  FIG. 4  is configured substantially similarly are shown and described in regard to  FIGS. 2-3 . In still various embodiments, the wall assembly  135  further includes a third wall  133  extended from the airfoil assembly  100 . In one embodiment, the third wall  133  is extended from a forward end corresponding to a leading edge  123  of the airfoil assembly  100 . In another embodiment, the third wall  133  may be extended from an aft end corresponding to a trailing edge  124  of the airfoil assembly  100 . In one embodiment, the first cavity  116  is defined between the third wall  133  and the first wall  131  extended between the airfoil assembly  100  and the hub  140 . 
     In still various embodiments, the third wall  133  may be extended from the airfoil assembly  100 , such as the base portion  110  thereof, within the passage  65  defined between the rotor assembly  90  and the first static assembly  310 . In another embodiment, the third wall  133  may be extended from an aft end of the rotor assembly  90 , such as to extend within the second passage  67  between the second static assembly and the aft side of the rotor assembly  90 . In various embodiments, the third wall  133  may define an opening  134  between the third wall  133  and the rotor assembly  90 . In one embodiment, the opening  134  between the third wall  133  and the rotor assembly  90  may be defined between the hub  140  of the rotor assembly  90  and the third wall  133 . In various embodiments, the third wall  133  extends radially inward toward the hub  140  to define the opening  134  between the third wall  133  and the rotor assembly  90  such as to admit the flow of cooling fluid therethrough to the airfoil assembly  100 . 
     In various embodiments, the base portion  110  defines a first inlet opening  111  in fluid communication with the first cavity  116 . In one embodiment, the first inlet opening  111  is defined through the forward end of the airfoil assembly  100  in fluid communication with the first cavity  116 . 
     Referring now to  FIGS. 5-7 , detailed exemplary embodiments of the airfoil assembly  100  are provided.  FIG. 5  provides a perspective view of an exemplary embodiment of the airfoil assembly  100 .  FIG. 6  provides a cross sectional view of the exemplary airfoil assembly  100  of  FIG. 5 .  FIG. 7  provides a top-down view of the exemplary embodiment of the airfoil assembly  100  provided in regard to  FIGS. 5-6 . Referring collectively to  FIGS. 5-7 , the airfoil assembly  100  defines a pressure side  121 , a suction side  122 , a leading edge  123 , and a trailing edge  124 . 
     Referring to  FIGS. 5-7 , in various embodiments, the base portion  110  of the airfoil assembly  100  includes a base portion wall  115  disposed within the base portion  110 . The base portion wall  115  defines a first plenum  113  and a second plenum  114  separated from one another by the base portion wall  115 . In one embodiment, the first plenum  113  in the base portion  110  is in fluid communication with the first cavity  116 . In another embodiment, the second plenum  114  in the base portion  110  is in fluid communication with the second cavity  117 . 
     In various embodiments, the airfoil assembly  100  further includes an airfoil structure  120  extended along the radial direction R from the base portion  110  and attached to the base portion  110 . For example, the airfoil structure  120  and the base portion  110  may be integrally formed together as the airfoil assembly  100  (e.g., casting, forging, machined, additive manufactured, etc., or combinations thereof). The airfoil assembly  100  defines a plurality of circuits  126 ,  127 ,  128 ,  129  in fluid communication with one or more of the first plenum  113  and the second plenum  114 . In various embodiments, the airfoil assembly  100  defines a first circuit  126  disposed in thermal communication at least at the leading edge  123  of the airfoil assembly  100 . In still various embodiments, the airfoil assembly  100  defines a second circuit  127  disposed in thermal communication at least at the trailing edge  124  of the airfoil assembly  100 . In another embodiment, the airfoil assembly  100  defines one or more of a third circuit  128  disposed between the first circuit  126  and the second circuit  127  along the chordwise direction  91 . It should be appreciated that in various embodiments, the airfoil assembly  100  may define a plurality of the first circuit  126 , the second circuit  127 , or the third circuit  128 . 
     In one embodiment, the airfoil assembly  100  defines the first circuit  126  in fluid communication with a first opening  101 . In another embodiment, the airfoil assembly  100  defines the second circuit  127  in fluid communication with a second opening  102 . The first circuit  126  and the second circuit  127  each extend at least partially through the airfoil structure  120 . 
     Referring still to  FIGS. 5-7 , in various embodiments, the airfoil assembly  100  further defines the third circuit  128  between the first circuit  126  and the second circuit  127  along the chordwise direction  91 . In still various embodiments, the third circuit  128  is in fluid communication with the first plenum  113 . In still yet various embodiments, the third circuit  128  defines a substantially serpentine passage or conduit through the airfoil structure  120 , such as to provide cooling between the leading edge  123  and the trailing edge  124  of the airfoil structure  120 . 
     In one embodiment, the first opening  101  may be disposed at the leading edge  123  of the airfoil structure  120 . In another embodiment, the second opening  102  may be disposed at the trailing edge  124  of the airfoil structure  120 . In still other embodiments, such as generally depicted in regard to  FIG. 5 , the airfoil structure  120  may define a third opening  103  through one or more of the pressure side  121 , the suction side  122 , a radially outward tip  125  ( FIG. 6 ), or combinations thereof, of the airfoil structure  120 . In various embodiments, one or more of the first circuit  126 , the second circuit  127 , or the third circuit  128  may be in fluid communication with the third opening  103 . 
     In various embodiments, the first circuit  126  may extend at the leading edge  123  of the airfoil assembly  100  and further fluidly couple to the second circuit  127  at the trailing edge  124 , the third circuit  128  between the leading edge  123  and the trailing edge  124 , or both, via a connecting circuit  129  ( FIGS. 8-12 ). The first circuit  126  may be in fluid communication with one or more of the first opening  101 , the second opening  102 , or the third opening  103 , or combinations thereof. In other embodiments, the second circuit  127  may extend at the trailing edge  124  of the airfoil assembly  100  and further fluidly couple to the first circuit  126  at the leading edge  123 , the third circuit  128  therebetween, or both, via the connecting circuit  129  ( FIGS. 8-12 ). The second circuit  127  may be in fluid communication with one or more of the first opening  101 , the second opening  102 , or the third opening  103 , or combinations thereof. 
     Referring now to  FIGS. 8-12 , schematic cross sectional views of the airfoil assembly  100  are generally provided. The embodiments provided in regard to  FIGS. 8-12  are configured substantially similarly as shown and described in regard to  FIGS. 1-7 . It should be appreciated that one or more walls, plenums, cavities, etc. such as generally depicted in regard to  FIG. 6  may be incorporated to define the plurality of circuits  126 ,  127 ,  128 ,  129  such as schematically depicted in regard to  FIGS. 8-12 . 
     Referring to  FIG. 8 , in one embodiment the first circuit  126  and the third circuit  127  are each in fluid communication with the first plenum  113 . The first plenum  113  receives the flow of first cooling fluid  83  from the first cavity  116  and first conduit  66 , such as described in regard to  FIGS. 2-4 . The embodiment provided in regard to  FIG. 8  may provide cooling to the leading edge  123  of the airfoil structure  120  via the first cooling fluid  83  defining a higher temperature and/or pressure relative to the second cooling fluid  84 . Additionally, the second circuit  127  is in fluid communication with the second plenum  114  to receive the flow of second cooling fluid  84  from the second cavity  117 . Additionally, or alternatively, the embodiment provided in regard to  FIG. 8  may provide cooling to the trailing edge  124  of the airfoil structure  120  via the second cooling fluid  84  defining a lower pressure and/or temperature relative to the first cooling fluid  83 . As yet another example, the embodiment provided in regard to  FIG. 8  may improve engine efficiency via reducing the amount of cooling flow extracted from a relatively higher pressure and higher temperature source, such as the first cooling fluid source  200  at the compressor exit  64  (e.g., temperature and pressure at the combustion section  26  at the compressor exit  64 ). 
     Referring now to  FIGS. 9-11 , in various embodiments the first circuit  126  and the second circuit  127  are each in fluid communication with the second plenum  114 . The first circuit  126  and the second circuit  127  are coupled together in fluid communication via a connecting circuit  129 . In one embodiment, the connecting circuit  129  extends across the chordwise direction  91  of the airfoil structure  120  to couple the first circuit  126  and the second circuit  127  in fluid communication. In various embodiments, the connecting circuit  129  is defined within the airfoil structure  120  to couple a plurality of chambers, cavities, etc. of a plurality of the first circuit  126 , the second circuit  127 , or the third circuit  128 . In one embodiment, the connecting circuit  129  is defined fluidly separate from the third circuit  128 , such as to provide the flow of second cooling fluid  84  to the leading edge  123  and the trailing edge  124  of the airfoil structure  120 . The third circuit  128  is in fluid communication with the first plenum  113 . In various embodiments, the third circuit  128  is fluidly separate or disconnected from the first circuit  126  and the second circuit  127  such as to provide the flow of first cooling fluid  83  through the airfoil structure  120  between the leading edge  123  and the trailing edge  124 . 
     Referring particularly to  FIG. 10 , in one embodiment, the connecting circuit  129  is defined at a radially inward or root portion of the airfoil assembly  100 . In one embodiment, the connecting circuit  129  is disposed in the base portion  110  of the airfoil assembly  100 . In various embodiments, the connecting circuit  129  is disposed in the airfoil structure  120  of the airfoil assembly  100 . In another embodiment, the airfoil structure  120  further includes a second connecting circuit  129 ( a ) defined at a radially outward or tip portion of the airfoil structure  120 . In various embodiments, the airfoil structure  120  may define one or more of the connecting circuits  129 ,  129 ( a ) disposed at a root portion, a tip portion, or radially therebetween through the airfoil structure  120 . 
     Referring to  FIGS. 9-10 , the second plenum  114  may be disposed forward (e.g., corresponding to the leading edge  123 ) within the airfoil assembly  100  and the first plenum  113  may be disposed aft (e.g., corresponding to the trailing edge  124 ) of the second plenum  114 , in which each plenum is separated by the base portion wall  115 . The flow of second cooling fluid  84  may be received at the second plenum  114  and routed aft through the airfoil assembly  100  from the first circuit  126 . The flow of second cooling fluid  84  may be received at the second plenum  114  and routed aft through the airfoil assembly  100  from the first circuit  126  to the second circuit  127 . 
     Referring to  FIG. 11 , the first plenum  113  may be disposed forward (e.g., corresponding to the leading edge  123 ) within the airfoil assembly  100  and the second plenum  114  may be disposed aft (e.g., corresponding to the trailing edge  124 ) of the first plenum  113 , in which each plenum is separated by the base portion wall  115 . The flow of second cooling fluid  84  may be received at the second plenum  114  and routed forward through the airfoil assembly  100  from the second circuit  127  to the first circuit  126 . 
     Referring to  FIGS. 9-11 , the flow of second cooling fluid  84  to the leading edge  123  and the trailing edge  124 , and the flow of first cooling fluid  83  therebetween along the chordwise direction  91 , enables providing a lower temperature and/or lower pressure source of cooling fluid to portions of the airfoil structure  120  that may be more prone to deterioration and damage due to combustion gases. Additionally, or alternatively, the lower temperature and/or lower pressure second cooling fluid  84  from the second cooling fluid source  300  enables reduced flow rates such as to reduce blockage at the exit of the compressor section  21  or at the combustion section  26 . 
     Referring to  FIG. 12 , in another embodiment the airfoil assembly  100  may include the first plenum  113  in the base portion  110  in fluid communication with the first cavity  116  and the second cavity  117  such as to define the first plenum  113  as a mixing chamber in fluid communication with the first cavity  116  and the second cavity  117 . The airfoil assembly  100  may further include the second plenum  114  in fluid communication with the first plenum  113 . In various embodiments, the base portion wall  115  may define one or more base portion apertures  118  through the base portion wall  115  such as to receive the combined flow of fluid  85  from the first plenum  113  into the second plenum  114 . The combined flow of fluid  85  includes the first cooling fluid  83  and the second cooling fluid  84  mixed at the first plenum  113  defining a mixing chamber. 
     In still various embodiments, the airfoil assembly  100  may include at the base portion  110  a mixer assembly  119  to promote mixing of the first cooling fluid  83  with the second cooling fluid  84 . For example, the mixer assembly  119  may define a swirler, a sparger device, a nozzle, etc. to condition the flows of fluid  83 ,  84  into the first plenum  113  defining a mixing chamber to promote mixing to provide the combined flow of fluid  85  to the second plenum  114 . The second plenum  114  may further be fluid communication with the first circuit  126 , the second circuit  127 , and the third circuit  128  to provide the combined flow of fluid  85  through the leading edge  123 , the trailing edge  124 , and portions therebetween of the airfoil structure  120 . 
     Portions of the engine  10 , such as the rotor assembly  90  and the airfoil assembly  100  depicted in regard to  FIGS. 1-12  and described herein, may be constructed as an assembly of various components that are mechanically joined or arranged such as to produce the embodiments of the rotor assembly  90  and the airfoil assembly  100  shown and described herein. The rotor assembly  90  and the airfoil assembly  100 , separately or together, may alternatively each or collectively be constructed as a single, unitary component and manufactured from any number of processes commonly known by one skilled in the art. For example, the rotor assembly  90  and the airfoil assembly  100  may be constructed as a single, unitary component. These manufacturing processes include, but are not limited to, those referred to as “additive manufacturing” or “3D printing”. Additionally, any number of casting, machining, welding, brazing, or sintering processes, or mechanical fasteners, or any combination thereof, may be utilized to construct the rotor assembly  90  and the airfoil assembly  100 . Furthermore, the rotor assembly  90  and the airfoil assembly  100  may be constructed of any suitable material for turbine engine rotor assemblies and airfoil assemblies, or more specifically high pressure or low pressure turbine rotor assemblies, including but not limited to, nickel- and cobalt-based alloys. Still further, flowpath surfaces and passages may include surface finishing or other manufacturing methods to reduce drag or otherwise promote fluid flow, such as, but not limited to, tumble finishing, barreling, rifling, polishing, or coating. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.