Patent Publication Number: US-2023146084-A1

Title: Gas turbine engine with clearance control system

Description:
GOVERNMENT SPONSORED RESEARCH 
     The project leading to this application has received funding from the European Union Clean Sky 2 research and innovation program under grant agreement No. CS2-ENG-GAM-2014-2015-01. 
    
    
     CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to Polish Patent Application No. P.439449, filed Nov. 5, 2021, which is a non-provisional application, and wherein the above application is hereby incorporated by reference in its entirety. 
     FIELD 
     The present subject matter relates particularly to gas turbine engines including clearance control structures. 
     BACKGROUND 
     Casings for gas turbine engines, such as turbine section casings surrounding turbine section rotors, generally require separable flanges and assembled casing and manifold portions due to internally and externally mounted components. Such components generally include brackets or hangers for turbine shrouds, or flanges for multiple casings. Additionally, since turbine casings surround turbine rotors, excessive deformation, thermal expansion or contraction, or bowing may result in excessive rub and undesired contact with the turbine rotors, which can result in loss in performance or operability. Certain casings may include assemblies via separable flanges to limit deformation or displacement during engine operation and thermal cycling. However, the inventors of the present disclosure have found that such designs require assembly and parts that add weight to the engine. Moreover, the inventors of the present disclosure have found that such designs may further inhibit the inclusion or placement of thermal control structures for more effective clearance control. 
     As such, the inventors of the present disclosure have found that there is a need for turbine casings that can overcome these limitations and provide improved thermal control, improved engine efficiency, and reduced weight. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present disclosure, 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 an exemplary schematic cross sectional view of an embodiment of a gas turbine engine in accordance with aspects of the present disclosure; 
         FIG.  2    is an exemplary schematic cross sectional view of an embodiment of a gas turbine engine in accordance with aspects of the present disclosure; 
         FIG.  3    is an exemplary schematic cross sectional view of an embodiment of a gas turbine engine in accordance with aspects of the present disclosure; 
         FIG.  4    is a schematic cross-sectional view of a portion of an embodiment of a gas turbine engine in accordance with aspects of the present disclosure; 
         FIG.  5    is a schematic cross-sectional view of a portion of an embodiment of a gas turbine engine in accordance with additional aspects of the present disclosure; 
         FIG.  6    is a perspective view of a portion of an embodiment of the gas turbine engine in accordance with aspects of the present disclosure; 
         FIGS.  7 A- 7 B  depict flowcharts outlining steps of a method for operating an engine in accordance with aspects of the present disclosure; 
         FIGS.  8 - 11    are exemplary schematic cross-sectional views of embodiments of a portion of a turbine section and casing in accordance with aspects of the present disclosure; 
         FIG.  12    is an exemplary perspective view of an embodiment of a portion of a manifold of the turbine section in accordance with aspects of the present disclosure; 
         FIGS.  13 A- 13 D  are exemplary sectional views of an embodiment of the manifold provided in  FIG.  12   ; 
         FIG.  14    is an exemplary schematic cross-sectional view of an embodiment of a portion of a turbine section and casing in accordance with aspects of the present disclosure; 
         FIG.  15    is an exemplary perspective view of an embodiment of a portion of a manifold of the turbine section in accordance with aspects of the present disclosure; 
         FIG.  16    is an exemplary schematic cross sectional view of an embodiment of a portion of a turbine section and casing in accordance with aspects of the present disclosure; 
         FIG.  17    is a detailed view of an exemplary schematic cross sectional view of the embodiment of  FIG.  16    in accordance with aspects of the present disclosure; 
         FIG.  18    is a top-down view of an exemplary embodiment of a plurality of pins of the thermal control ring in accordance with aspects of the present disclosure; 
         FIG.  19    is an exemplary schematic of flows of air through the turbine section and casing of  FIG.  16    in accordance with aspects of the present disclosure; 
         FIG.  20    is a perspective view of a portion of the engine in accordance with aspects of the present disclosure; and 
         FIG.  21    is a cross-sectional view of the embodiment of the engine provided in  FIG.  20    in accordance with aspects of the present disclosure. 
     
    
    
     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 disclosure. 
     DETAILED DESCRIPTION 
     Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure. 
     The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary. 
     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. 
     The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. 
     The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein. 
     Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin. 
     Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. 
     Pressure values, and ranges thereof, are in absolute pressure measurement (psia) or equivalent. Values and ranges of pressure provided herein may be converted to ranges in gauge pressure, or other pressure units, or other units, measurements, or combinations thereof that correspond to the values and/or ranges disclosed herein. 
     The term “overall power output” refers to a maximum rated power output of an engine. 
     The term “operating envelope” refers to a cycle, mission, or set of maneuvers at which the engine may normally operate. In one embodiment, a landing-takeoff (LTO) cycle may define an operating envelope. The LTO cycle including one or more combinations of startup, idle, takeoff, cruise, and approach engine operating conditions may collectively define the operating envelope. In various embodiments, the cruise condition defines a majority of the operating envelope, such as to define a majority of an operating time or duration of the engine operation. In certain embodiments, the cruise condition is between approximately 55% and 75% of the operating envelope. Stated differently, the cruise condition may define approximately 55% to approximately 75% of the duration of engine operation from startup to shutdown following approach operating condition. In another embodiment, the cruise condition may define approximately 60% to approximately 70% of the duration of engine operation. 
     The term “cruise operating condition” may further refer to mid-power engine operating condition. The term “takeoff operating condition” may refer to a full power condition and “idle operating condition” may refer to a low power condition, and “cruise operating condition” is a power or thrust condition therebetween. In some embodiments, the cruise condition corresponds to approximately 75% to approximately 90% of an overall power output of the engine. In still certain embodiments, the cruise condition corresponds to approximately 80% to 88% of the overall power output of the engine. 
     A “third stream” as used herein means a non-primary air stream capable of increasing fluid energy to produce a minority of total propulsion system thrust. A pressure ratio of the third stream may be higher than that of the primary propulsion stream (e.g., a bypass or propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of an airflow through the third stream with a primary propulsion stream or a core air stream, e.g., into a common nozzle. 
     In certain exemplary embodiments an operating temperature of the airflow through the third stream may be less than a maximum compressor discharge temperature for the engine, and more specifically may be less than 350 degrees Fahrenheit (such as less than 300 degrees Fahrenheit, such as less than 250 degrees Fahrenheit, such as less than 200 degrees Fahrenheit, and at least as great as an ambient temperature). In certain exemplary embodiments these operating temperatures may facilitate heat transfer to or from the airflow through the third stream and a separate fluid stream. Further, in certain exemplary embodiments, the airflow through the third stream may contribute less than 50% of the total engine thrust (and at least, e.g., 2% of the total engine thrust) at a takeoff condition, or more particularly while operating at a rated takeoff power at sea level, static flight speed, 86 degree Fahrenheit ambient temperature operating conditions. 
     Furthermore in certain exemplary embodiments, aspects of the airflow through the third stream (e.g., airstream, mixing, or exhaust properties), and thereby the aforementioned exemplary percent contribution to total thrust, may passively adjust during engine operation or be modified purposefully through use of engine control features (such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features) to adjust or optimize overall system performance across a broad range of potential operating conditions. 
     The term “turbomachine” or “turbomachinery” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output. 
     The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc. 
     The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof. 
     The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc. each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high turbine” or “high speed turbine” at the engine. 
     The term “at,” as used herein to refer to a location of a first object relative to a second object (e.g., the first object located or positioned at the second object) refers to the first object being positioned wholly or partially within the second object, the first object contacting the second object, or the first object being positioned closest to the second object (relative to any other surrounding relevant components). 
     One or more components of the turbomachine engine described herein below may be manufactured or formed using any suitable process, such as an additive manufacturing process, such as a 3-D printing process. The use of such a process may allow such component to be formed integrally, as a single monolithic component, or as any suitable number of sub-components. In particular, the additive manufacturing process may allow such component to be integrally formed and include a variety of features not possible when using prior manufacturing methods. For example, the additive manufacturing methods described herein may allow for the manufacture of passages, conduits, cavities, openings, casings, manifolds, double-walls, heat exchangers, or other components, or particular positionings and integrations of such components, having unique features, configurations, thicknesses, materials, densities, fluid passageways, headers, and mounting structures that may not have been possible or practical using prior manufacturing methods. Some of these features are described herein. 
     Suitable additive manufacturing techniques in accordance with the present disclosure include, for example, Fused Deposition Modeling (FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjets, laser jets, and binder jets, Stereolithography (SLA), Direct Selective Laser Sintering (DSLS), Electron Beam Sintering (EBS), Electron Beam Melting (EBM), Laser Engineered Net Shaping (LENS), Laser Net Shape Manufacturing (LNSM), Direct Metal Deposition (DMD), Digital Light Processing (DLP), Direct Selective Laser Melting (DSLM), Selective Laser Melting (SLM), Direct Metal Laser Melting (DMLM), and other known processes. 
     Suitable powder materials for the manufacture of the structures provided herein as integral, unitary, structures include metallic alloy, polymer, or ceramic powders. Exemplary metallic powder materials are stainless steel alloys, cobalt-chrome, aluminum alloys, titanium alloys, nickel based superalloys, and cobalt based superalloys. In addition, suitable alloys may include those that have been engineered to have good oxidation resistance, known as “superalloys” which have acceptable strength at the elevated temperatures of operation in a gas turbine engine, e.g. Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-850, ECY 768, 282, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys. The manufactured objects of the present disclosure may be formed with one or more selected crystalline microstructures, such as directionally solidified (“DS”) or single-crystal (“SX”). 
     Embodiments of a gas turbine engine including an improved clearance control system are provided. The engine reduces weight and tubes, manifolds, or conduits outside of an outer core casing or fan casing by reducing or eliminating air extracted from a fan bypass passage for cooling at a turbine section. Embodiments provided herein allow for engines without fan casings, such as open rotor engines or propfan engines, to have and operate improved clearance control, cooling systems, or air systems for turbine sections and/or bearing assemblies. It should be appreciated that while such embodiments may be applied to turbofan engines including nacelles and fan casings, embodiments provided herein allow for engines without nacelles, fan casings, or other structures surrounding the fan section to receive air for turbine section cooling, clearance control, or bearing assemblies. 
     The improved gas turbine engine provided herein may additionally, or alternatively, allow for lower-pressure and/or lower-temperature air to be removed from the compressor section for cooling or clearance control at the turbine section and bearing assembly. Certain clearance control systems may generally utilize high-energy air (i.e., high-pressure and/or high-temperature air), such as from aft stages of a high pressure compressor, and mix with one or more other sources of air, such as from other compressor stages or from the fan air stream. Such high-energy air reduces engine efficiency, such as by removing energy from the thermodynamic and combustion process, or by requiring greater reduction in heat load before the air is appropriate for cooling or clearance control at the turbine section. Still further, certain clearance control systems may not be suitable for additionally providing air to a bearing assembly for cooling, buffer air, or other uses at the bearing assembly. 
     Another aspect of the disclosure is directed to an improved turbine casing allowing for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. Embodiments of an engine, casing, and manifold provided herein include integral, unitary structures such as may be formed by additive manufacturing processes that would not have heretofore been possible or practicable. Embodiments depicted and described herein allow for improved and advantageous positioning of thermal control rings for improved clearance control response, improved formation and positioning of openings, passages, and conduits to allow for more efficient heat transfer fluid utilization and movement, and reduced weight, such as via obviating flanges and sub-assemblies into integral components. Particular combinations of these features allow for improved heat transfer properties and reduced thermal gradients. Improved heat transfer properties particularly include a lower heat transfer coefficient at certain features, such as at the plurality of walls that form thermal control rings as provided herein. Such improvements may mitigate or eliminate undesired or excessive deformation, ovalization, bowing, or other changes in casing geometry that may adversely affect deflections or result in undesired contact to the turbine rotors. 
     Embodiments provided herein include, e.g., an integral, unitary high speed turbine casing and turbine center frame or mid-turbine frame positioned downstream of the high speed turbine and upstream of a low- or intermediate-pressure turbine. Embodiments provided herein further include, e.g., an integral, unitary clearance control manifold configured to provide heat transfer fluid to thermal control rings. The integral, unitary structures may further allow for improved positioning of the thermal control rings relative to the turbine rotors, such as to provide improved clearance control across the turbine rotor assembly. 
     As used herein, the term “integral, unitary” as used to describe a structure refers to the structure being formed integrally of a continuous material or group of materials with no seams, connections joints, or the like. The integral, unitary structures described herein may be formed through additive manufacturing to have the described structure, or alternatively through a casting process, etc. 
     Referring now to the drawings,  FIG.  1    is a schematic cross-sectional view of an exemplary gas turbine engine  10  herein referred to as “engine  10 ” as may incorporate various embodiments of the present disclosure. Particular embodiments of the engine  10  may be configured as a turbofan, turboprop, turboshaft, or propfan gas turbine engine, or one or more gas turbine engines configured as hybrid-electric gas turbine engines, or other gas turbine engine configuration. 
     As shown in  FIG.  1   , the engine  10  has a longitudinal or axial centerline axis  12  that extends therethrough for reference purposes parallel to an axial direction A. In general, the engine  10  may include a turbomachine  14  disposed downstream from a fan section  16 . 
     The engine  10  includes a compressor section  21  in serial flow arrangement with a turbine section  27 . The turbomachine  14  may generally include a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  may be formed from multiple casings. The outer casing  18  encases, in serial flow arrangement, the compressor section  21 , a combustion section  26 , and the turbine section  27 . In a particular embodiment, the compressor section  21  includes a booster or low speed compressor  22  and a high speed compressor  24 . In a still particular embodiment, the turbine section  27  includes a first turbine assembly or high speed turbine  28  and a second turbine assembly or low speed turbine  30  (e.g., including vanes  116  and rotor blades  118 ). A jet exhaust nozzle section  32  is positioned downstream of the turbine section  27 . A high speed shaft or spool  34  drivingly connects the high speed turbine  28  to the high speed compressor  24 . A low speed shaft or spool  36  drivingly connects the low speed turbine  30  to the low speed compressor  22 . The low speed spool  36  may also be connected to a fan shaft or spool  38  of the fan section  16 . In particular embodiments, the low speed spool  36  may be connected directly to the fan spool  38  such as in a direct-drive configuration. In alternative configurations, as is depicted in phantom in  FIG.  1   , the low speed spool  36  may be connected to the fan spool  38  via a gear assembly  37 , such as to configure the engine  10  as an indirect-drive or geared-drive configuration allowing for a higher or lower rotational speed of the fan spool  38  versus the low speed spool  36 . Such gear assemblies may be included between any suitable shafts/spools within engine  10  as desired or required. 
     Although depicted and described as a two-spool engine including the high speed spool  34  separately rotatable from the low speed spool  36 , it should be appreciated that the engine  10  may be configured as a three-spool engine including the high speed spool  34 , the low speed spool  36 , and a third spool or intermediate speed spool positioned in serial flow arrangement between the high speed spool  34  and the low speed spool  36 . Accordingly, the compressor section  21  may include an intermediate speed compressor separately rotatable from the high speed compressor  24  and the low speed compressor  22 . Similarly, the turbine section  27  may include a third turbine assembly or an intermediate speed turbine separately rotatable from the high speed turbine  28  and the low speed turbine  30 . The intermediate speed compressor and the intermediate speed turbine may together be coupled to form an intermediate speed spool fluidly between the high speed spool and the low speed spool. 
     It should further be appreciated that in certain embodiments the low speed turbine  30  or second turbine assembly described herein generally refers to a separately rotatable spool downstream of the high speed turbine or first turbine assembly. As such, the second turbine assembly may include an intermediate speed turbine or a low speed turbine positioned aft or downstream of the high speed turbine. 
     As shown in  FIG.  1   , the fan section  16  includes one or more axially-spaced stages of a plurality of fan blades  40  that are coupled to and that extend radially outwardly from the fan spool  38 . An annular fan casing or nacelle  42  circumferentially surrounds the fan section  16  and/or at least a portion of the turbomachine  14 . It should be appreciated that for the embodiment depicted the nacelle  42  is supported relative to the turbomachine  14  by a plurality of circumferentially-spaced outlet guide vanes  44 . 
     A bypass airflow passage  48  is formed downstream of one or more stages of the plurality of fan blades  40  and around an outer portion of the turbomachine  14 . In a particular embodiment, such as depicted in  FIG.  1   , the bypass airflow passage  48  is defined at a downstream section  46  of the nacelle  42  (downstream of the outlet guide vanes  44 ) and between the nacelle  42  and the outer portion of the turbomachine  14 . 
     However, in other embodiments, it should be appreciated that the low speed compressor  22  may form one or more stages of the fan section  16 , such as depicted in  FIG.  3   . As such, the bypass airflow passage  48  may generally include any flowpath downstream of one or more stages of the plurality of fan blades  40  or the low speed compressor  22  and bypassing or surrounding at least a portion of the high speed compressor  24 , and having a flow of bypass air  177  therethrough provide thrust. Accordingly, certain embodiments of the engine  10  provided herein may be configured as a third stream or adaptive cycle engine having a plurality of bypass airflow passages  48  downstream of one or more stages of the plurality of fan blades  40  and/or the low speed compressor  22  and upstream of at least a portion of the high speed compressor  24 , with one or more of which configured as a “third stream.” 
     The engine  10  includes a computing system  1210  configured to perform operations. The computing system  1210  is communicatively coupled to the turbomachine  14  and/or a starter motor (not depicted) to adjust, modulate, maintain, change, or articulate any one or more control surfaces to generate the flows of air, one or more embodiments of the flow of heat transfer fluid, and/or a liquid and/or gaseous fuel in accordance with aspects of the present disclosure provided herein. The computing system  1210  can generally correspond to any suitable processor-based device, including one or more computing devices. Certain embodiments of the computing system  1210  include a full authority digital engine controller (FADEC), a digital engine controller (DEC), or other appropriate computing device configured to operate the engine  10 . 
     The computing system  1210  may include one or more processors  1212  and one or more associated memory devices  1214  configured to perform a variety of computer-implemented functions, such as steps of the methods described herein. As used herein, the term “processor” refers not only to integrated circuits referred to in the art as being included in a computer, but also refers to a controller, microcontroller, a microcomputer, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), a Field Programmable Gate Array (FPGA), and other programmable circuits. Additionally, the memory  1214  can generally include memory element(s) including, but not limited to, computer readable medium (e.g., random access memory (RAM)), computer readable non-volatile medium (e.g., flash memory), a compact disc-read only memory (CD-ROM), a magneto-optical disk (MOD), a digital versatile disc (DVD), non-transitory computer-readable media, and/or other suitable memory elements or combinations thereof. 
     The computing system  1210  may include control logic  1216  stored in the memory  1214 . The control logic  1216  may include computer-readable instructions that, when executed by the one or more processors  1212 , cause the one or more processors  1212  to perform operations, such as outlined in one or more steps of the method  1000  provided further below. In still various embodiments, the memory  1214  may store charts, tables, functions, look ups, schedules etc. corresponding to the flows, or rates, pressures, or temperatures associated with the flows of air, heat transfer fluid, or fuel provided herein. The instructions can be software written in any suitable programming language or can be implemented in hardware. Additionally, and/or alternatively, the instructions can be executed in logically and/or virtually separate threads on the processor(s). 
     The computing system  1210  may also include a communications interface module  1230 . In various embodiments, the communications interface module  1230  can include associated electronic circuitry that is used to send and receive data. As such, the communications interface module  1230  of the computing system  1210  can be used to receive data from one or more control surfaces, sensors, measurement devices, or instrumentation, or calculations or measurements corresponding to one or more portions of the engine  10  provided herein, and may execute one or more steps of the method  1000  provided herein. The computing system(s)  1210  can also include a network interface used to communicate, for example, with the other components of engine  10 . The network interface can include any suitable components for interfacing with one or more network(s), including for example, transmitters, receivers, ports, controllers, antennas, and/or other suitable components. 
     It should be appreciated that the communications interface module  1230  can be any combination of suitable wired and/or wireless communications interfaces and, thus, can be communicatively coupled to one or more components of the apparatus via a wired and/or wireless connection. As such, the computing system  1210  may obtain, determine, store, generate, transmit, or operate any one or more steps of the method described herein via a distributed network. For instance, the network can include a SATCOM network, ACARS network, ARINC network, SITA network, AVICOM network, a VHF network, a HF network, a Wi-Fi network, a WiMAX network, a gatelink network, etc. 
     Referring now to  FIG.  2   , an exemplary embodiment of an open rotor configuration of the engine  10  depicted and described with regard to  FIG.  1    is provided. The embodiment of the engine  10  provided in  FIG.  2    is configured substantially similarly as provided in  FIG.  1   . However, in  FIG.  2   , the open rotor configuration of the engine  10  does not have a fan casing or the nacelle  42  (depicted in  FIG.  1   ) surrounding the plurality of fan blades  40 . The bypass airflow passage  48  is formed downstream of the plurality of fan blades  40 , or particularly downstream of the outlet guide vanes  44 , and radially outward of the outer portion of the turbomachine  14 . 
     Referring now to  FIG.  3   , an exemplary embodiment of an open rotor configuration in accordance with  FIG.  2    is provided. The embodiment provided in  FIG.  3    further includes a plurality of bypass airflow passages  48  formed downstream of the plurality of fan blades  40 , such as described above. In the particular embodiment depicted, the engine  10  includes a first bypass airflow passage  48 A and a second bypass airflow passage  48 B. The second bypass airflow passage  48 B is extended from a location between the low speed compressor  22  and the high speed compressor  24  to an exhaust to atmosphere (although in other embodiments the second bypass airflow passage  48 B may extend to the first bypass airflow passage  48 A). An articulating vane or door structure  43  may be positioned at the second bypass airflow passage  48 B. The door structure  43  may include any appropriate type of actuatable wall, vane, door, or other structure configured to desirably alter a flow of air  172  received from a core gas flowpath  70  and allowed through the second bypass airflow passage  48 B, such as depicted schematically via arrows  177 . The second bypass airflow passage  48 B may be referred to as a third stream. 
     Although  FIG.  3    depicts the engine  10  having a three-stream or adaptive cycle with an open rotor configuration, it should be appreciated that the adaptive cycle configuration may also include a nacelle surrounding the fan section, such as depicted and described with regard to  FIG.  1   . In such a manner, it should further be appreciated that although certain advantages and benefits provided herein may provide benefits for turbofan engines having nacelles, embodiments and arrangements of the components provided herein may overcome issues or challenges that are particular to open rotor configurations. 
     Referring now to  FIGS.  4 - 5   , enlarged cross-sectioned views of an engine  10  configured in a similar manner as one or more of the exemplary engines  10  depicted in  FIGS.  1 - 3    are provided.  FIGS.  4 - 5    depict walled conduits, manifolds, tubes, or other structures forming flowpaths configured to extract or receive a flow of air, depicted schematically via arrows  91 , from the compressor section  21  and provide the flow of air  91  to the turbine section  27 . The flow of air  91  provided to the turbine section  27  may be utilized for cooling blades, vanes, shrouds, or other portions of the turbine section  27 . In certain embodiments, the turbine section  27  includes a turbine frame  308  positioned in serial flow arrangement between the first turbine assembly or the high speed turbine  28  and the second turbine assembly or low speed turbine  30 . In still particular embodiments, a bearing assembly  200  is included at the turbine frame  308 . Accordingly, the turbine frame  308  may provide a static mount or support structure at which the bearing assembly  200  is positioned to support rotation of one or more spools (e.g., low speed spool  36  or high speed spool  34 ). The turbine frame  308  further includes any appropriate quantity of conduits, manifolds, or passages  309 , or other structures for allowing at least a portion of the flow of air  91  (e.g., depicted further below as flow of air  193 ) to the bearing assembly  200 . The flow of air to the bearing assembly  200  may provide cooling or buffer air at the bearing assembly  200 , such as to attenuate vibrations from the spool or generate desired bearing or rotor clearances. In other embodiments, the flow of air  91  is provided to the gear assembly  37  positioned at the fan section  16 , to the compressor section  21 , to the turbine section  27 , or to the jet exhaust nozzle section  32 . 
     The engine  10  includes a first conduit  110  extended in fluid communication from the compressor section  21  to the turbine section  27 . The first conduit  110  is configured to communicate the flow of air  91  from the compressor section  21  to a first location  271  at the turbine section  27 . The first conduit  110  forms a flow passage separate from the core gas flowpath  70 . In a particular embodiment, the first conduit  110  provides the flow of air  91  from the compressor section  21  to the turbine section  27  while bypassing the combustion section  26 . 
     A first heat exchanger  141  is positioned in thermal communication with the flow of air  91  through the first conduit  110 . The first heat exchanger  141  is configured to receive heat or thermal energy from the flow of air  91  through the first conduit  110 . Accordingly, the first heat exchanger  141  is configured to cool the flow of air  91  through the first conduit  110  before the flow of air  91  is provided to the turbine section  27 . The first heat exchanger  141  is configured as any appropriate heat exchanger for extracting heat or thermal energy from the flow of air  91  and receiving or transmitting heat or thermal energy to a heat transfer fluid, depicted schematically via arrows  221 . Particular embodiments of the engine  10  may include a fluid system  220  configured to flow the heat transfer fluid  221  as a lubricant, a liquid and/or gaseous fuel, a hydraulic fluid, a supercritical fluid, a refrigerant, or an appropriately cooler air or inert gas. The fluid system  220  provides the heat transfer fluid  221  into thermal communication with the flow of air  91  via the first heat exchanger  141 . In a particular embodiment depicted in  FIG.  9    (discussed in more detail below), the heat transfer fluid  221  is a liquid fuel provided to the combustion section  26 . However, it should be appreciated that the heat transfer fluid  221  may be provided and utilized in any appropriate way, including, but not limited to, as a lubricant for a bearing system, an anti-icing fluid, fuel, or actuation fluid. 
     Referring still to  FIGS.  4 - 5   , the engine  10  includes a second conduit  120  extended from the first conduit  110  downstream of the first heat exchanger  141  (relative to the flow of air  91  from the compressor section  21  to the turbine section  27 ). The second conduit  120  is extended in fluid communication to a second location  272  at the turbine section  27 . A flow control device  130  is positioned at the second conduit  120 . The flow control device  130  is configured to selectively adjust, alter, modulate, or otherwise change an amount of the flow of air  91  from the first conduit  110  through the second conduit  120 . 
     In various embodiments, the second conduit  120  includes an inlet portion  121  and an outlet portion  122 . The inlet portion  121  is fluidly coupled to the first conduit  110  and the flow control device  130 . The inlet portion  121  extends from the first conduit  110  to provide a portion of the flow of air  91 , depicted schematically via arrows  192 , to the flow control device  130 . The outlet portion  122  is fluidly coupled to the flow control device  130  and the second location  272  of the turbine section  27 . The outlet portion  122  extends from the flow control device  130  to provide at least a portion of the flow of air  192  to the second location  272  at the turbine section  27 . In such a manner, it will be appreciated that for the embodiment depicted, the flow control device  130  is positioned between the inlet and outlet portions  121 ,  122  of the second conduit  120 . 
     The flow control device  130  may be a valve or any appropriate device for regulating, directing, controlling, or otherwise modulating an amount of flow of fluid across a passage or flowpath. The flow control device  130  may include an actuated valve or an automatic valve driven by an electric energy source, a pneumatic energy source (e.g., air, or particularly, at least a portion of the flow of air  91 ), or a fluid source (e.g., liquid and/or gaseous fuel, hydraulic fluid, lubricant, or combinations thereof). The flow control device  130  may include ball valves, shuttle valves, or other appropriate type of valve or flow regulating device in accordance with the embodiments depicted and described herein. Accordingly, the flow control device  130  is configured to modulate the amount of flow of fluid through the outlet portion  122  of the second conduit  120 , such as depicted schematically via arrows  94 . 
     In a particular embodiment, the engine  10  includes a third conduit  123  extending from the flow control device  130  to a third location  273  at the turbine section  27 , in fluid communication with both the flow control device  130  and the third location  273 . The flow control device  130  may therefore be a three-way valve configured to selectively change the amount of the flow of air  91  from the first conduit  110  through the inlet portion  121  of the second conduit  120  to one or both of the third conduit  123  and the outlet portion  122  of the second conduit  120 . Accordingly, the flow control device  130  may be configured to modulate an amount of the flow of air  192  through the outlet portion  122  of the second conduit  120 , such as depicted schematically via arrows  194 , and furthermore modulate and egress of at least a portion of the flow of air  192  through the third conduit  123 , such as depicted schematically via arrows  195 . The third conduit  123  may form a bypass passage to further allow for selective adjustment, control, or modulation of the flows of air through the flow control device  130 . In a particular embodiment, the third conduit  123  allows for a portion of the air extracted from the first conduit  110  to bypass the outlet portion  122  of the second conduit  120  and egress to the third location  273  at the turbine section  27 . In certain embodiments, the third location  273  allows for bypassing a clearance control system  275  (described below) and allowing the flow of air  195  to enter the turbine section  27  at the core gas flowpath  70  downstream of the clearance control system  275 , or to mix with the flow of air  193  at the turbine frame  308 , or to vent to ambient (not depicted). 
     Referring still to  FIGS.  4 - 5   , as briefly noted above, the turbine section  27  includes the clearance control system  275 . Exemplary embodiments of improved clearance control systems are depicted in  FIGS.  8 - 16   , including a casing  300 , manifold assemblies  316 , and thermal control rings  314  such as provided therein. However, it should be appreciated that the clearance control system  275  depicted in  FIGS.  4 - 5    may include any appropriate structure or assembly for controlling, adjusting, or otherwise modulating a dimension between a rotor blade tip and a surrounding shroud or wall at the turbine section  27 , otherwise referred to as tip clearance. The clearance control system  275  may be an active clearance control (ACC) system configured to dynamically control tip clearance. Particularly, the ACC system may be configured to desirably modulate the tip clearance based on an engine operating condition via the flow of air  94  received from the second conduit  120  and provided to a surrounding shroud at the turbine section  27 . The volumetric or mass flow rate of the flow of air  94  is regulated or modulated by the flow control device  130 . Modulating the amount of the flow of air  94  to the clearance control system  275  allows the tip clearance to be desirably regulated across various engine operating conditions and associated changes in temperature at the turbine section  27 . As temperatures and rotor speeds change at the turbine section  27  across various engine operating conditions, the flow control device  130  modulates the amount of the flow of air  94  provided to clearance control system  275  to maintain or provide a desired tip clearance. With regard to a landing-takeoff cycle (LTO) of the engine  10  and an aircraft, engine operating conditions include startup, idle, takeoff, climb, cruise, approach, or reverse thrust. However, it should be appreciated that other engine operating conditions and cycles may be applicable. 
     Referring still to  FIGS.  4 - 5   , the second location  272  at the turbine section  27  is at the clearance control system  275 . Accordingly, the second conduit  120 , or particularly the outlet portion  122  of the second conduit  120 , is fluidly coupled to the turbine section  27  to provide the flow of air  94  to the clearance control system  275  such as described herein. In a particular embodiment, the clearance control system  275  is operably coupled to a first turbine assembly or the high speed turbine  28  at the turbine section  27 . Accordingly, the engine  10  is configured to receive the flow of air  91  from the compressor section  21  and provide the portion of the flow of air  94  (from the flow of air  91 ) to the clearance control system  275  at the high speed turbine  28  via the second conduit  120 . 
     In still particular embodiments, the first conduit  110  is fluidly coupled to the turbine frame  308  positioned between the first turbine assembly, or the high speed turbine  28 , and a second turbine assembly, or low speed turbine  30 . The turbine frame  308  may include a plurality of vanes  310  in circumferential arrangement and positioned between the turbines  28 ,  30 . The first location  271  at the turbine section  27  is at the turbine frame  308 . Accordingly, in such embodiments, the first conduit  110  is configured to provide at least a portion of the flow of air  91  to the turbine frame  308  at the first location  271 . In a particular embodiment, schematic arrows  193  depict a portion of the flow of air at the first conduit  110  downstream of a juncture with the second conduit  120 . The flow of air  193  is provided to the turbine frame  308  via the first conduit  110 . In particular embodiments further depicted and described with regard to  FIGS.  8 - 16   , the flow of air  193  may be provided to the casing  300  and through the plurality of vanes  310  at the turbine frame  308 , such as depicted schematically via arrows  99 . 
     Referring to  FIGS.  4 - 5   , the turbine frame  308  may include or form one or more passages  309  configured to provide fluid communication of the flow of air  193  to the bearing assembly  200 . The flow of air  193  may provide a buffer fluid for operation of the bearing assembly  200 . The buffer fluid may desirably control or attenuate vibrations, or allow or generate desired clearances or vibratory responses at the bearing assembly  200  or the rotors to which the bearing assembly is coupled. 
     Referring now specifically to  FIG.  5   , in a particular embodiment, the engine  10  includes a second heat exchanger  142  in thermal communication with a flow of air at the bypass airflow passage  48 . The second heat exchanger  142  may be configured as a surface heat exchanger configured to receive heat or thermal energy from the flow of air  194  downstream of the flow control device  130  at the second conduit  120 . The heat transfer fluid at the second heat exchanger  142  is a flow of air through the bypass airflow passage  48  of the engine  10 , such as depicted schematically via arrows  177 . The second heat exchanger  142  configured as a surface heat exchanger has a heat exchange surface at the bypass airflow passage  48  and is configured to place the flow of air  194  at the second conduit  120  in thermal communication with the flow of bypass air  177  at the bypass airflow passage  48 . In a particular embodiment, the second heat exchanger  142  is positioned at the outlet portion  122  of the second conduit  120  and upstream of the second location  272  at the turbine section  27 . 
     Referring back generally to both  FIGS.  4 - 5   , in a particular embodiment, the first conduit  110  includes an inlet manifold  111  configured to receive the flow of air  91  from a circumferential compressor location  211  at the compressor section  21 . It should be appreciated that although the embodiments depicted in  FIGS.  4 - 5    depict a single circumferential compressor location  211 , the inlet manifold may be configured to receive the flow of air  91  from a plurality of circumferential compressor locations  211 . 
     Referring now to  FIG.  6   , a perspective view of an embodiment of a portion of an engine  10  in accordance with one or more of  FIGS.  1  through  3    is provided. The embodiment provided in  FIG.  6    may be configured substantially similarly as described in regard to the embodiments in  FIGS.  4 - 5   . In  FIG.  6   , the engine  10  may include a plurality of inlet manifolds  111  evenly-spaced or asymmetrically-spaced along the circumferential direction C around the compressor section  21 . In various embodiments, the plurality of inlet manifolds  111  includes two (2) or more inlet manifolds. In one embodiment, the plurality of inlet manifolds  111  includes three (3) inlet manifolds. In another embodiment, the plurality of inlet manifolds  111  includes four (4) inlet manifolds and up to 30 inlet manifolds  111 . 
     In  FIG.  6   , the first conduit  110  includes a collector  115  configured to receive the flow of air  91  from the inlet manifold  111 . In particular embodiments, the plurality of inlet manifolds  111  is fluidly coupled to a single collector  115  to provide a collected or unified flow of air  91  to the first heat exchanger  141 . The collector  115  may provide the flow of air  91  to the first heat exchanger  141 , such as described herein. 
     In a still particular embodiment, the first conduit  110  includes an outlet manifold  112  configured to fluidly communicate the flow of air  91  from the first heat exchanger  141  to the turbine section  27  at the first turbine location  271  at the turbine section  27 . The engine  10  may include a plurality of outlet manifolds  112  evenly-spaced or asymmetrically-spaced along the circumferential direction C around the turbine section  27 . In various embodiments, the plurality of outlet manifolds  112  includes two (2) or more outlet manifolds. In one embodiment, the plurality of outlet manifolds  112  includes three (3) outlet manifolds. In another embodiment, the plurality of outlet manifolds  112  includes four (4) outlet manifolds and up to 30 outlet manifolds. In various embodiments, the second conduit  120  is extended in fluid communication from one or more of the plurality of outlet manifolds  112  of the first conduit  110 . The plurality of outlet manifolds  112  may accordingly extend to a plurality of first turbine locations  271  at different circumferential positions at the turbine section  27 . 
     It should be appreciated that although the embodiments depicted in  FIGS.  4 - 5    depict a single circumferential first turbine location  271 , the first turbine location  271  may include a plurality of circumferential first turbine locations  271 . 
     Embodiments of the engine  10  provided in  FIGS.  4 - 5    may include the first conduit  110  as a fixed area flowpath from the compressor section  21  to the turbine section  27 . Stated differently, the first conduit  110  may include various cross-sectional areas or convergent and divergent flowpaths. However, the first conduit  110  and the circumferential compressor location  211  may define fixed or non-articulatable flowpath areas. Such fixed area flowpath allows for a constant volumetric or mass flow rate of the flow of air  91  from the compressor section  21  through the first conduit  110  with respect to a corresponding engine operating condition. Stated differently, the fixed area flowpath allows for the first conduit  110  to receive a corresponding flow rate of the flow of air  91  relative to the particular engine operating condition. Accordingly, embodiments of the engine  10  provided herein allow for constant flows of air  91  in thermal communication with the flow of heat transfer fluid  221  at the first heat exchanger  141 . For instance, flow rates of the heat transfer fluid  221 , such as a fuel flow rate or lubricant flow rate, may be controlled via a schedule, table, graph, or curve indicative of the flow rate versus the engine operating condition. In one embodiment, the flow of air  91  at the first conduit  110  may generally be fixed as a ratio or proportion of the overall flow of air entering the core engine inlet  20  into the compressor section  21 . In another embodiment, the flow of air  91  at the first conduit  110  may generally be fixed as a ratio or proportion of the flow of air entering the high speed compressor  24  from the low speed compressor  22 . 
     The engine  10  may particularly include a variable area flowpath at the second conduit  120  via the flow control device  130 . Accordingly, the engine  10  may allow a fixed flow of air  193  to the turbine frame  308 , such as for the bearing assembly  200 , and a variable flow of air  194  to the clearance control system  275 . The flow control device  130  may adjust, articulate, or otherwise modulate the flow of air  194  to the clearance control system  275  as a function of engine operating condition. Modulation of the flow of air  194  via the flow control device  130  may be a function of inlet air speed (into the turbomachine  14  via an inlet  20 ), or inlet air pressure (e.g., corresponding to altitude of the engine  10  during operation or at one or more engine operating conditions described above), or inlet air temperature, or combinations thereof. Modulation of the flow of air  194  via the flow control device  130  may additionally, or alternatively, be a function tip clearance at the turbine section  27 , or a predetermined schedule corresponding to wear or deterioration at the turbine section  27 . 
     Certain embodiments of the engine  10  include particular placements of the circumferential compressor location  211  at particular axial stages or other location at the compressor section  21  corresponding to particular pressure ranges of the flow of air  91  during operation of the engine  10 . In various embodiments, the circumferential compressor location  211  from which the flow of air  91  is received from the core gas flowpath  70  corresponds to a compressor location having an airflow therethrough at a pressure between approximately 20 pounds per square inch (psi) and approximately 60 psi during an engine operating condition corresponding to between approximately 55% and approximately 75% of an operating envelope. In another embodiment, the circumferential compressor location  211  from which the flow of air  91  is received from the core gas flowpath  70  may corresponding to a compressor location having an airflow therethrough at a pressure between approximately 30 pounds per square inch (psi) and approximately 50 psi during the engine operating conditions such as described herein. 
     Accordingly, embodiments of the engine  10  provided herein allow for the clearance control system  275  and the bearing assembly  200  to operate and receive air from the compressor section  21 . In certain embodiments, the engine  10  provided herein allows for the clearance control system  275  to receive the flow of  91  from the compressor section  21  rather than from the bypass airflow passage  48 . Furthermore, or alternatively, the engine  10  provided herein allows for the flow of air  91  to be received from upstream, forward, or lower-pressure stages of the compressor section  21  in contrast to other compressor bleed systems that may receive high energy air from downstream, aft, or higher-pressure stages of a compressor section. Certain of these other compressor bleed systems may further mix the higher-energy air with lower-energy (i.e., lower pressure, lower temperature, or both) corresponding to the bypass airflow passage. Still further, or alternatively, embodiments of the engine  10  provided herein allow for a constant flow of air  91  through the first conduit  110  to maintain purge and backflow margin at the turbine frame  308  and bearing assembly  200 . 
     Referring now to  FIGS.  7 A- 7 B , a flowchart outlining steps of the method  1000  for operating an engine is provided. The steps of the method  1000  may be stored as instructions and/or executed as operations by embodiments of the engine  10  and the computing system  1210  provided herein. Accordingly, the method  1000  may be a computer-implemented method in which one or more steps is stored as instructions at the memory  1214  at the computing system  1210  and/or executed by one or more processors  1212  at the computing system  1210 . The computing system  1210  may cause embodiments of the engine such as described herein with regard to  FIGS.  1 - 6    to perform operations such as outlined in the flowchart in  FIGS.  7 A- 7 B  and described further herein with regard to method  1000 . 
     Referring to the flowchart in  FIGS.  7 A- 7 B , and in conjunction with any one or more embodiments depicted in  FIGS.  1 - 6   , the method  1000  includes at  1010  initiating rotation of one or both of a high speed spool or a low speed spool to, e.g., generate compressed air for combustion within a combustion section of a core engine. In various embodiments, a motive force, such as a starter motor or turbine air starter (not shown), initiates rotation of one or both of the high speed spool  34  or the low speed spool  36  to generate an initial airflow through the core gas flowpath  70  into the combustion section  26  for mixing with a liquid and/or gaseous fuel before igniting to generate combustion gases. 
     The method  1000  further includes at  1020  compressing a flow of air through the compressor section. During operation of the engine  10 , a flow of air  171  is received at the fan section  16 . A portion of the flow of air  171  enters the turbomachine  14  through the core engine inlet  20 , such as depicted schematically via arrows  172 . The flow of air  172  is pressurized across successive rows or stages of compressor blades at the compressor section  21 . Particularly, the low speed compressor  22  may include a low pressure compressor or booster relative to the high speed compressor  24  including a high pressure compressor. In certain embodiments, a portion of the flow of air  172  compressed by the low speed compressor  22  may be bled or re-routed from the core gas flowpath  70 , such as to control stall, surge, or operability at one or both of the compressors  22 ,  24 . The high speed compressor  24  receives the flow of air  172  and further compresses the flow of air, such as depicted schematically via arrows  173  in  FIGS.  1 - 3   . The successive stages of compressor blades energize the flow of air  173 , such as to increase the pressure and temperature of the flow of air  173  before entering the combustion section  26 , such as depicted via arrows  174 . 
     The method  1000  includes at  1030  extracting a portion of the compressed flow of air from the compressor section, such as described above. The method  1000  at  1030  may particularly include extracting the portion of compressed flow of air into a first conduit and bypassing a combustion section, such as provided above with regard to the first conduit  110 . The method  1000  includes at  1040  flowing the extracted portion of the compressed flow of air through the first conduit (e.g., first conduit  110 ) to a turbine section. In a particular embodiment, the first conduit bypasses the combustion section when flowing the extracted portion of compressed flow of air to the turbine section. With regard to  FIGS.  1 - 6   , a portion of the flow of air at the compressor section  21  is bled or removed from the core gas flowpath  70  and provided to the first conduit  110 , such as depicted schematically via arrows  91  in  FIGS.  1 - 5   . Particular embodiments depicted herein may receive the flow of air  91  from the compressed flow of air  173 ,  174  from the high speed compressor  24 . In still other embodiments, the flow of air  91  may be received from the compressed flow of air  172  from the low speed compressor  22 . 
     It should be appreciated that embodiments of the engine  10  provided herein advantageously receive relatively lower-pressure and lower-temperature flows of air from the compressor section  21 , and may further avoid structures, complexities, actuatable devices, valves, and associated weight and efficiency losses related to mixing high-pressure and high-temperature air with low-pressure and low-temperature air from the fan bypass airflow passage. It should furthermore be appreciated that, while particular operating conditions and operating envelopes are provided herein, the engine  10  and/or method  1000  provided herein allows for performing one or more steps at any engine operating condition, including up to 100% of an overall power output. However, particular advantages and benefits are provided herein with regard to operation of the engine at engine operating conditions defining a majority of an operating envelope. As such, methods and structures provided herein allow for improved efficiency and reduced fuel consumption. 
     In various embodiments, the method  1000  at  1030  includes extracting the portion of the compressed flow of air when the compressed flow of air at the compressor section is between approximately 20 pounds per square inch (psi) and approximately 60 psi. In a particular embodiment, the method  1000  at  1030  includes extracting the portion of the compressed flow of air when the compressed flow of air at the compressor section is between approximately 30 psi and approximately 50 psi. In a particular embodiment, the method  1000  includes at  1035  receiving the portion of the compressed flow of air from the compressor section, in which the portion of the compressed flow of air is between approximately 20 psi and approximately 60 psi, or between approximately 30 psi and approximately 50 psi. In a still particular embodiment, the method  1000  at  1030  and/or  1035  is performed continuously or constantly relative to a discrete engine operating condition, such as to allow for a fixed flow of air relative to the discrete engine operating condition. 
     In a still particular embodiment, the method  1000  includes at  1028  operating the engine at an engine condition corresponding to between approximately 55% and approximately 75% of an operating envelope, or between approximately 60% and approximately 70% of the operating envelope, such as described above. In certain embodiments, one or both steps of the method  1000  at  1030  and at  1035  is preceded by, or contemporaneous to, the method  1000  at  1028 . In still certain embodiments, the method  1000  includes at  1029  operating the engine between approximately 75% and approximately 90% of the overall power output (e.g., rated thrust) of the engine, such as described above. In a still particular embodiment, the method  1000  at  1029  includes operating the engine between approximately 80% and approximately 88% of the overall power output of the engine. In certain embodiments, one or more ranges provided herein may define a discrete engine operating condition at which the method  1000  at  1030  and/or  1035  is performed continuously or constantly. In still particular embodiments, the method  1000  includes performing the steps at  1028  and  1029  concurrently. 
     The method  1000  may include at  1050  flowing, via a fluid system, a heat transfer fluid in thermal communication with the extracted portion of compressed flow of air, such as described above. In a particular embodiment, the fluid system  220  depicted in  FIGS.  4 - 5    is a liquid and/or gaseous fuel system configured to provide a flow of liquid and/or gaseous fuel to the compressed flow of air  174  to generate combustion gases  175 . In such an embodiment, the fuel is the heat transfer fluid  221  in thermal communication with the flow of air  91  via the first heat exchanger  141 . The flow of fuel receives heat or thermal energy from the relatively hotter flow of air  91 , which may advantageously alter certain properties of the fuel, such as viscosity, density, or other property that may desirably affect combustion, fuel-air mixing, swirl, emissions generation, vibrations, or smoke and particulate generation. 
     In certain embodiments, the method  1000  may further include flowing, via the fluid system, a plurality of heat transfer fluids in thermal communication with the extracted portion of compressed flow of air. In various embodiments, the method  1000  includes providing one or more flows of fuel, lubricant, hydraulic fluid, refrigerant, a supercritical fluid, or another flow of air at the heat transfer fluid in thermal communication with the extracted flow of air. 
     The method  1000  may further include modulating the flow of the heat transfer fluid to control a temperature of the extracted flow of air (e.g., flow of air  91 ). Modulating the flow of heat transfer fluid may include adjusting a mass or volumetric flow rate, pressure, or temperature of the heat transfer fluid provided in thermal communication with the extracted flow of air. 
     As provided above, the flow of liquid and/or gaseous fuel is mixed with the compressed air from the compressor section and ignited to form combustion gases  175 . The combustion gases  175  flow from the combustion section  26  to the turbine section  27 , and particularly to the high speed turbine  28  and the low speed turbine  30 . As the combustion gases  175  expand at the turbine section  27 , energy is released to drive rotation of the respective turbines  28 ,  30 , which drives their respective spools  34 ,  36 , compressors  22 ,  24 , and fan blades  40 . 
     It should be appreciated that the combustion gases  175  release variable amounts of heat at the turbine section  27  based on the engine operating condition. Accordingly, heat release and turbine rotor speed may alter the tip clearance between turbine rotor blade tips and surrounding shrouds, such as further described below. It should be appreciated that improved aerodynamic and operating efficiencies are generally achieved by minimizing tip clearances. Accordingly, clearance control systems are utilized to modulate the tip clearance based on engine operating condition to improve engine efficiency and performance. 
     The method  1000  may further include at  1060  selectively flowing a portion of the flow of air through a second conduit (e.g., second conduit  120 ) extended from the first conduit (e.g., first conduit  110 ) downstream of the heat exchanger (e.g., first heat exchanger  141 ). In a particular embodiment, the method  1000  includes at  1062  varying or modulating, via a flow control device (e.g., flow control device  130 ) at the second conduit extended from the first conduit, the portion of the flow of air extracted to the second conduit (e.g., second conduit  120 ) from the first conduit (e.g., first conduit  110 ) downstream of the heat exchanger (e.g., first heat exchanger  141 ). In a still particular embodiment, the method  1000  includes at  1063  modulating, via the flow control device, a second portion of the flow of air extracted from the first conduit to the third conduit extended from the flow control device, such as depicted in  FIGS.  4 - 5    via arrows  195 . In a still particular embodiment, the method  1000  at  1060  is executed contemporaneously with the method  1000  at one or more of steps  1028 ,  1030 , or  1035 . Accordingly, the method  1000  may allow for continuous, constant, or fixed flow of air from the compressor section through the first conduit, while modulating or varying the flow of air through the second conduit. In particular embodiments, the method  1000  allows for continuous, constant, of fixed flow of air from the compressor section through the first conduit and to the turbine section, or particularly the bearing assembly, while modulating or varying the flow of air through the second conduit to a clearance control system. As such, modulating the flow of air through the second conduit allows for a variable flow of air through to a clearance control system (e.g., clearance control system  275 ) independent of whether the operating condition of the engine is steady-state (e.g., non-transient or non-varying) or transient (e.g., changing). 
     The method  1000  may further include at  1070  selectively varying, altering, or modulating a tip clearance at a clearance control system based on the flow of air received from the second conduit via step  1060  and/or  1062 . It should be appreciated that the method  1000  provided herein may further provide for a method for operating a clearance control system and bearing assembly. Such methods may allow for variable flow rate, temperature, pressure, or other physical property of the flow of air through the second conduit to the clearance control system, while allowing for substantially constant or continuous flows of air through the first conduit relative to an engine operating condition. 
     Although not depicted in  FIGS.  7 A- 7 B , the method  1000  may further include generating a flow of bypass air through a bypass airflow passage. A portion of the flow of air  171  passes across the plurality of fan blades  40  and bypasses the turbomachine  14 , such as depicted via arrows  176  in  FIGS.  1 - 3   . The flow of air  176  that enters the bypass airflow passage  48 , depicted schematically via arrows  177 , is large in volume or mass and cold relative to the flow of air pressurized by the compressor section  21  within the turbomachine  14 .  FIG.  5   , which may be applied to the embodiments of the engine  10  in any of  FIGS.  1 - 3   , particularly depicts the relatively cold flow of bypass air  177  in thermal communication with the flow of air  194  via the second heat exchanger  142 . Accordingly, the method  1000  may further include at  1064  thermally communicating, via the second heat exchanger, the flow of bypass air with the portion of the flow of air extracted to the second conduit. 
     The embodiment of the engine  10  depicted and described with regard to  FIG.  5    may allow for increased magnitudes of heat transfer from the flow of air  194 , such as via the flow of bypass air  177  at the bypass airflow passage  48 . Furthermore, the embodiment depicted in  FIG.  5   , when applied to an open rotor configuration such as depicted in  FIG.  2   , may overcome challenges associated with removing nacelles and passages, tubes, or conduits that may route through nacelles to provide air for heat exchangers, clearance control systems, and/or bearing assemblies. Accordingly, the method  1000 , when applied to an open rotor configuration such as described herein, may provide for a method for operating an open rotor engine, or particularly, a method for operating a clearance control system for an open rotor engine, or more particularly, a method for operating a clearance control system and bearing assembly for an open rotor engine. 
     Referring now to  FIG.  8   , an enlarged cross sectioned view is provided of a turbine section portion of a turbomachine  14  in accordance with one or more of  FIGS.  1 - 3   , as may incorporate various embodiments of the present disclosure. As shown in  FIG.  8   , a first turbine assembly is formed by the high speed turbine  28 . A first stage  50  of the first turbine assembly includes a plurality of first turbine rotor blades  58  extended within the core gas flowpath  70 , and further includes an annular array of stator vanes  54  (only one shown) axially spaced from an annular array of turbine rotor blades  58  (only one shown) at the high speed turbine  28 . In a particular embodiment, the high speed turbine  28  further includes a last stage  60  which includes an annular array of stator vanes  64  (only one shown) axially spaced from an annular array of turbine rotor blades  68  (only one shown). The turbine rotor blades  58 ,  68  extend radially outwardly from and are coupled to the high speed spool  34  ( FIG.  1   ,  FIG.  2   ). The stator vanes  54 ,  64  and the turbine rotor blades  58 ,  68  at least partially define the core gas flowpath  70  for routing combustion gases from the combustion section  26  ( FIG.  1   ,  FIG.  2   ) through the high speed turbine  28 . 
     As further shown in  FIG.  8   , the high speed turbine  28  may include one or more shroud assemblies, each of which forms an annular ring about an annular array of rotor blades. For example, a shroud assembly  72  may form an annular ring around the annular array of rotor blades  58  of the first stage  50  and the annular array of turbine rotor blades  68  of the last stage  60 . In general, the shroud assembly  72  is radially spaced from blade tips  76 ,  78  of each of the rotor blades  58 ,  68 . A radial or clearance gap CL is defined between the blade tips  76 ,  78  and respective inner surfaces of the shroud segments  77 . The shroud assembly  72  generally reduces leakage from the core gas flowpath  70 . The shroud assembly  72  can include a plurality of walls forming thermal control rings  314  that assist in controlling thermal growth of the shroud thereby controlling the radial deflection or clearance gap CL. Thermal growth in the shroud assemblies is actively controlled with the clearance control system  275 . The clearance control system  275  is used to minimize radial blade tip clearance CL between the outer blade tip and the shroud, particularly during cruise operation of the engine, such as described herein. 
     Downstream along the core gas flowpath  70 , or aft of the high speed turbine  28 , is a second turbine assembly formed by the low speed turbine  30 . As previously described herein, the second turbine assembly is rotatably separate from the first turbine assembly, such as described in regard to the high speed turbine  28  and the low speed turbine  30  above with reference to  FIG.  1   . 
     The casing  300  surrounds the high speed turbine  28 . The casing  300  includes a plurality of vanes  310  extended through the core gas flowpath  70  aft of the first turbine assembly formed by the high speed turbine  28  and forward of the second turbine assembly formed by the low speed turbine  30 . The shroud assembly  72  is coupled to the casing  300  at an outer casing wall  312 . The outer casing wall  312  is an annular wall surrounding the shroud assembly  72  and extended along a circumferential direction C relative to the centerline axis  12  ( FIGS.  1 - 3   ). The outer casing wall  312  is extended along the axial direction A forward of the rotor blades  58  of the first stage  50  of the high speed turbine  28  (also referred to as the first stage of rotor blades  58 ) and aft of the rotor blades  68  of the second or last stage  60  of the high speed turbine  28  (also referred to as the second stage of rotor blades  68 ). 
     The plurality of vanes  310  is extended from the outer casing wall  312 . The plurality of vanes  310  is extended into the core gas flowpath  70 , In certain embodiments further described herein, one or more of the plurality of vanes  310  may be hollow or include conduits or passages allowing for fluid flow within the vane. The outer casing wall  312  of the casing  300  is extended along the axial direction A from a downstream end or trailing edge of the aft-most stage of the rotor blades  68  to at least an upstream end or leading edge of the plurality of vanes  310 , such as depicted at dimension B in  FIG.  8   . 
     It should be appreciated that conventional turbine casings include separable or joined flanges, such as bolted flanges or welded flanges, between a high speed turbine casing and a downstream casing, such as an inter-turbine frame, mid-turbine frame, intermediate speed turbine casing, or low speed turbine casing. Embodiments of the casing  300  provided herein, include unitary, integral structures, such as formed by one or more additive manufacturing processes. Embodiments provided herein further form integral, continuous, compliant structures, allowing for the unitary, integral extension of the casing  300  such as provided herein, or further including one or more features integrally formed to the casing  300  such as provided herein. 
     A plurality of walls forming thermal control rings  314  is extended along the circumferential direction C and extended outward along a radial direction R from the outer casing wall  312 . In various embodiments, the thermal control rings  314  include forward thermal control rings  3141  positioned outward along the radial direction R from the first stage of rotor blades  58 , or particularly from the blade tips  76  of the rotor blades  58 , of the high speed turbine  28 . In certain embodiments, such as depicted in  FIG.  8   , the forward thermal control rings  3141  are positioned in alignment along the axial direction A to the first stage of rotor blades  58  (overlapping axial positions). In another particular embodiment, the thermal control rings  314  include aft thermal control rings  3142  positioned outward along the radial direction R from the last stage  60  of rotor blades  68 , or particularly from the blade tips  78  of the rotor blades  68 , of the high speed turbine  28 . In certain embodiments, such as depicted in  FIG.  8   , the aft thermal control rings  3142  are positioned in alignment along the axial direction A to the last stage  60  of rotor blades  68  of the high speed turbine  28  (overlapping axial positions). 
     The forward and aft thermal control rings  3141  and  3142  are provided to more effectively control blade tip clearance CL (shown in  FIG.  8   ) with a minimal amount of time lag and thermal control airflow (cooling or heating depending on operating conditions). The forward and aft thermal control rings  3141  and  3142  are formed with the outer casing wall  312  as an integral, singular, unitary structure of the casing  300 . The thermal control rings  314  provide thermal control mass to more effectively move the shroud segments  77  along the radial direction R to adjust the blade tip clearances CL. Such clearance control may provide for lower operational specific fuel consumption (SFC). 
     The integral, unitary structure of the thermal control rings  314  and the outer casing wall  312 , with the outer casing wall particularly extended aft of the second or last stage of the rotor blades  68  of the high speed turbine  28 , may allow for improved clearance control, improved thermal control, and improved cooling flow. The structures provided herein allow for the thermal control rings  314  to be positioned radially outward of and in axial alignment with each stage of the high speed turbine rotor, such as to improve clearance control at each respective stage. The structures provided herein further allow for obviating flanges between the high speed turbine and an intermediate turbine frame between the high speed turbine and a downstream low speed turbine (or intermediate speed turbine, such as described herein). 
     Embodiments of the integral casing provided herein are generally produced by one or more additive manufacturing processes such as described above. Although additive manufacturing may generally be applied to form various structures or integrate various components, it should be appreciated that combinations of integrated structures provided herein may overcome issues associated with integrating structures while providing unexpected benefits. In one instance, axially-extended casings may generally be susceptible to thermal distortion that may ovalize the core flowpath, which may adversely affect rotor operation as the rotors may rub within a non-concentric flowpath. As such, simple integration of relatively hot casings surrounding the high speed turbine with the relatively cooler casing surrounding downstream vanes proximate to the low speed turbine may adversely affect overall engine operation. In another instance, such large, axially-extended masses may require additional cooling flow, which results in increased fuel consumption and overall losses in engine performance. 
     Embodiments of the engine provided herein overcome such issues at least in part by the positioning of the thermal control rings in axial alignment and radially outward of the respective stages of the high speed turbine blades. Removing flanges between a casing surrounding the high speed turbine rotors and a vane casing or frame downstream of the high speed turbine allows for the thermal control rings to be advantageously positioned as disclosed herein. 
     Other embodiments of the engine provided herein overcome such issues at least in part by improved cooling flow structures, passages, and conduits. In various embodiments, a manifold assembly  316  surrounds the thermal control rings  314  along the circumferential direction C and the axial direction A. The manifold assembly  316  is configured to provide a flow of fluid, such as the flow of air  192  from the compressor section  21  such as depicted and described in regard to  FIGS.  4 - 5   , to the thermal control rings  314 . 
     Referring still to  FIG.  8    and now also to  FIGS.  9 - 11   , and  FIG.  14   , further exemplary embodiments are provided. The embodiment depicted in  FIG.  8   ,  FIG.  9   , and  FIG.  14    may be configured similarly as one another, such as further described below.  FIGS.  9 - 11    provide views of flows of fluid and openings at various cross-sections of an embodiment of the engine  10  at different circumferential positions of the engine  10 . Each of the embodiments may be formed via one or more manufacturing methods known in the art. In  FIG.  14   , the embodiment provided may include double-wall structures that may be formed via an additive manufacturing process. Various embodiments provided herein may be formed as integral, unitary structures, such as via an additive manufacturing process or other appropriate manufacturing process. 
     Referring to the various embodiments depicted in  FIGS.  8 - 11    and  FIG.  14   , the manifold assembly  316  is extended along the axial direction A forward and aft of the plurality of axially-spaced stages of the plurality of walls forming the thermal control rings  314 . In a particular embodiment, such as depicted in  FIG.  14   , the manifold assembly  316  is extended aft along the axial direction A of the plurality of vanes  310 . In various embodiments, such as in the exemplary embodiment of  FIG.  8   , the manifold assembly  316 , the outer casing wall  312 , and the plurality of walls forming the thermal control rings  314  of the casing  300  is a single, integral, unitary structure, such as described herein. In particular embodiments, such as in the exemplary embodiment of  FIG.  8   , the manifold assembly  316  includes a plurality of concentric walls integrally formed and surrounding the outer casing wall  312 . In certain embodiments, the manifold assembly  316  includes an inner manifold  1316  radially inward of and concentric to an outer manifold  2316 . In still certain embodiments, the inner manifold  1316  is a double wall structure concentric to the outer manifold  2316 . 
     Referring particularly to  FIGS.  9 - 10   , certain embodiments of the casing  300  include a corrugated feature  399 . The corrugated feature  399  includes a shape defining ridges or grooves configured to mitigate formation of thermal expansion stresses at the casing  300 . In certain embodiments, the corrugated feature  399  is formed at the manifold assembly  316 . In a still particular embodiment, the corrugated feature  399  may be formed at an inner manifold  1316  or an outer manifold  2316 . The corrugated feature  399  may allow for the unitary, integral formation of the manifold assembly  316  with the outer casing wall  312 , such as described in various embodiments herein. 
     Referring now briefly to  FIG.  15   , the manifold assembly  316  includes a plurality of openings  318  surrounding the plurality of walls forming the thermal control rings  314  at the casing  300 . The plurality of openings  318  allow for the flow of fluid, depicted schematically via arrows  91 , to come into thermal communication with the thermal control rings  314  for desired heat transfer effect. In various embodiments, the plurality of openings  318  include an inlet opening  3181  configured to allow the flow of air  91  into a first cavity  1321  in thermal communication with the thermal control rings  314 , as described further below. The plurality of openings  318  may further include an outlet opening  3182  configured to allow at least a portion of the flow of air  91 , depicted schematically via flow of air  92 , to egress the first cavity  1321  and enter an inner wall conduit  1326  such as described further below. 
     An inlet opening wall  381  is extended between an outer portion  346  and an inner portion  347  of the double wall structure formed by the inner manifold  1316 . The inlet opening wall  381  forms an inlet opening flowpath  382  that extends through the double wall structure fluidly separated from the inner wall conduit  1326 . The inlet opening  3181  and the inlet opening wall  381  allow for the flow of air  91  to pass from a conduit  1324  surrounding the inner manifold  1316  to enter a plenum  383  formed between adjacent thermal control rings  314 . Particularly, the inlet opening wall  381  extends between the outer portion  346  and inner portion  347  of the inner manifold  1316 . The inlet opening flowpath  382  formed by the inlet opening wall  381  allows the flow of air  91  to enter the plenum  383  while being fluidly segregated from the flow of air  92  through the inner wall conduit  1326 . 
     Referring back particularly to  FIGS.  9 - 10   , as discussed, the manifold assembly  316  includes the inner manifold  1316  surrounding the thermal control rings  314  along the circumferential direction C and the axial direction A. The manifold assembly  316  depicted further includes the outer manifold  2316  surrounding the inner manifold  1316 , as discussed above. A passage wall  1318  is extended to the outer manifold  2316  from the inner manifold  1316  to form a passage  1320  within the passage wall  1318 . 
     In certain embodiments, such as depicted in  FIG.  8   , the outer manifold  2316  of the manifold assembly  316  is extended along the axial direction A at or aft the plurality of vanes  310 . The outer manifold  2316  is further connected to the outer casing wall  312  at or aft of the plurality of vanes  310 . In still certain embodiments, such as depicted in  FIGS.  9 - 11   , the inner manifold  1316  is extended to a location forward along the axial direction A of the plurality of vanes  310  (terminating forward of the plurality of vanes  310 ). The inner manifold  1316  is also extended to a location aft along the axial direction A of the plurality of walls forming the thermal control rings  314 . As such, the inner manifold  1316  is connected to the outer casing wall  312  forward of the plurality of vanes  310  and aft of the thermal control rings  314 . 
     The first cavity  1321  discussed above with reference to  FIG.  15    (also depicted in  FIGS.  9 - 11   ) is formed between the inner manifold  1316  and the outer casing wall  312 . The thermal control rings  314  are surrounded by the inner manifold  1316  at a location within the first cavity  1321  between the inner manifold  1316  and the outer casing wall  312 . The passage  1320  allows for fluid communication with the first cavity  1321  between the inner manifold  1316  and the outer casing wall  312 . The passage  1320  further allows for the flow of air  91  to enter into thermal communication with the thermal control rings  314 . 
     In various embodiments, the conduit  1324  briefly mentioned above is formed between the outer manifold  2316  and the inner manifold  1316 . The conduit  1324  is in fluid communication with the first cavity  1321  and is fluidly separated from passage  1320  by the passage wall  1318 . In particular embodiments, the passage wall  1318  is extended from the outer manifold  2316  to the inner manifold  1316  through the conduit  1324 . 
     Referring particularly to  FIGS.  9 - 11   , and further in regard to  FIG.  14   , the conduit  1324  is further extended in fluid communication through one or more of the plurality of vanes  310 .  FIG.  10    and  FIG.  14    particularly depict the flow of air  91  entering into thermal communication and fluid communication with the thermal control rings  314  in the first cavity  1321 .  FIG.  10    particularly depicts the flow of air  91  entering into thermal communication and fluid communication with the thermal control rings  314  in the first cavity  1321 . In various embodiments, the first cavity  1321  is formed to direct the flow of fluid to thermal contact portions of the thermal control rings directly, such as in a perpendicular direction.  FIG.  11    and  FIG.  14    particularly depict the flow of air  92  egressing from the first cavity  1321  through the conduit  1324  and then in serial flow through one or more of the plurality of vanes  310  (as airflow  99 , discussed below). In certain embodiments, the thermal control rings  314  are formed with the outer casing wall  312  to desirably improve clearance control. In one embodiment, such as depicted in  FIG.  13 B , the thermal control ring  314  includes outer surfaces extended as a ridge, groove, or at acute or zig-zagging angles (see more detailed description below). 
     Referring briefly particularly to  FIG.  14   , and further depicted in the detailed perspective view in  FIG.  15   , in certain embodiments, the inner manifold  1316  is a double wall structure forming the inner wall conduit  1326  between the double wall structure of the inner manifold  1316 . The inner wall conduit  1326  may extend in fluid communication to a second cavity  1322  formed between the outer casing wall  312  and an outer wall  170  of the core gas flowpath  70 . In such embodiments, the unitary, integral casing  300 , or furthermore integral to embodiments of the manifold assembly  316 , allow for separate flows into the plurality of vanes  310 . Particularly, the flow of air  91  enters the conduit  1324  from a compressor section such as depicted and described with regard to  FIGS.  1 - 6   . A portion of the flow of air  91 , depicted via arrows  92 , flows into the first cavity  1321  and then into the inner wall conduit  1326  formed at the double wall structure. The flow of air  92  then flows into one or more of the plurality of vanes  310 . Furthermore, another portion of the flow of air  91 , depicted via arrows  99 , remains in the conduit  1324  and flows into one or more of the plurality of vanes  310 . In certain embodiments, the flows  92 ,  99  are isolated or fluidly separated from one another until mixing at the plurality of vanes  310 . In other embodiments, the flows  92 ,  99  remain fluidly separated and are provided to separate respective vanes  310 , or separate conduits within each vane  310 . Embodiments of the casing  300  and the manifold assembly  316  allow for improved thermal efficiency and improved overall engine efficiency, such as via providing secondary uses of the flow of fluid after thermal communication with the thermal control rings  314 , rather than outputting the flows to atmosphere or to an under-cowl area of the engine. 
     In certain embodiments, the outer wall  170  of the core gas flowpath  70  forms the outer shroud segment  77  of the shroud assembly  72 . The outer shroud segment  77  is exposed to the core gas flowpath  70 , and may include thermal barrier coatings or materials configured to withstand heat from the combustion gases. The outer shroud segment  77  may further be configured to at least partially rub with one or more stages of blades at the core gas flowpath  70 . 
     Referring still to  FIG.  14   , and further depicted in  FIG.  15   ,  FIG.  16    providing a side view of the casing  300  of  FIG.  15   , and  FIG.  17    providing a close-up view of Section A in  FIG.  16   , the inner manifold  1316  includes a plenum wall  1319  extended from the inner manifold  1316  and surrounding the thermal control ring  314 . In certain embodiments, the plenum wall  1319  is extended radially inward from the inner portion  347  of the inner manifold  1316 . The plenum wall  1319  may be formed as an integral, unitary, or monolithic structure with the inner manifold  1316  including the outer portion  346  and the inner portion  347 . The first cavity  1321  is formed between an outer surface  1325  of the thermal control ring  314  and the plenum wall  1319 . 
     Referring particularly to  FIGS.  16  and  17   , the thermal control ring  314  includes a wall or body  332  extended outward, such as outward along the radial direction R, from the outer casing wall  312 . In various embodiments, such as described above with regard to the plurality of thermal control rings  314 , the body  332  is extended substantially annularly along the circumferential direction C ( FIGS.  1 - 3   ). 
     Referring more particularly to  FIG.  17   , the body  332  forms an internal flowpath  330  to allow a flow of fluid through the thermal control ring  314 . The flow of fluid through the body  332  allows for a temperature or thermal gradient at the thermal control ring  314  to be desirably controlled, altered, or modulated by changes in temperature or flow rate of the flow of fluid through the flowpath  330  at the body  332 . The flow of fluid through the body  332  may furthermore allow for one or more structures attached or integrally formed to the thermal control ring  314 , such as the outer casing wall  312  or the shroud assembly  72 , to move based at least in part on thermal changes provided by the flow of fluid, such as to desirably control the clearance gap CL ( FIG.  8   ) between the rotor blades  58 ,  68  and the shroud assembly  72 . 
     Referring still to  FIG.  17   , the exemplary casing  300  depicted further includes a plurality of pins  334  extended along a radial direction R of the engine  10  incorporating the casing  300  from the outer casing wall  312  to the body  332 . Referring briefly also to  FIG.  18   , a top-down view of the plurality of pins  334  depicts each pin  334  is depicted. As shown in  FIGS.  17  and  18   , each pin  334  spaced apart from one another along an axial direction A of the engine  10  incorporating the casing  300  and along a circumferential direction C of the engine  10  incorporating the casing  300  ( FIG.  18   ). In such a manner, adjacent pins  334  define a space  336  therebetween. 
     Referring back particularly to  FIG.  17   , the flowpath  330  extended radially through the body  332  is further extended in fluid communication to the gap or space  336  provided between the plurality of pins  334 . The thermal control ring  314  may form the flowpath  330  as a plurality of discrete, round or slotted flowpaths in adjacent arrangement along the circumferential direction C. In other embodiments, the thermal control ring  314  forms the flowpath  330  as a plurality of arcuate sections extended at least partially along the circumferential direction C. The flow of air, depicted schematically via arrows  91  is received and provided in fluid communication with the thermal control rings  314  in accordance with any one or more embodiments depicted and described above with regard to  FIGS.  1 - 15   . 
     During operation, the flow of air  91  passes through the spaces  336  and across the plurality of pins  334  to enter into the flowpath  330  within the body  332 . During operation, the flow of air  91  progresses radially through the body  332  and egresses the body  332  through an outlet opening  338  at the flowpath  330 . The outlet opening  338  is formed by the body  332  distal to the spaces  336  to allow for fluid communication from the flowpath  330  to the inner wall conduit  1326  formed within the double wall structure of the inner manifold  1316 . The flow of fluid egressed from the thermal control ring  314 , depicted schematically via arrows  92 , may flow through the inner wall conduit  1326  in accordance with any one or more embodiments depicted and described with regard to  FIGS.  1 - 15   . 
     Referring still to  FIG.  17   , in various embodiments, a seal  1323  is positioned to contact the outer surface  1325  of the thermal control ring  314  and the plenum wall  1319 . Additionally, or alternatively, the seal  1323  may be formed or positioned in contact with the inner portion  347  of the inner manifold  1316  and the outer surface  1325  of the body  332  of the thermal control ring  314 . The seal  1323  inhibits a flow of fluid through the first cavity  1321 . In a particular embodiment, the seal  1323  may form a structural member configured to provide structural support to the inner manifold  1316  and/or the thermal control ring  314 . The seal  1323  may further support the body  332  relative to the plurality of pins  334 . In certain embodiments, the seal  1323  is a braze, weld, or other member attaching the plenum wall  1319  to the thermal control ring  314  at the first cavity  1321 . It should be appreciated that the seal  1323  and the plenum wall  1319  may each extend substantially co-directional with the thermal control ring  314  as either a monolithic annular component or as a plurality of arcuate sections positioned in circumferential arrangement. 
     In particular embodiments, the outer casing wall  312 , the plurality of pins  334 , and the body  332  of the thermal control rings  314  are a unitary, integral structure, such as may be formed by an additive manufacturing process, or other appropriate manufacturing process. In still particular embodiments, the inner portion  347 , the outer portion  346 , and the plenum wall  1319  are together formed as a unitary, integral structure of the inner manifold  1316 . In certain embodiments, the thermal control rings  314  and outer casing wall  312  are a unitary structure separate from the inner manifold  1316 . In still certain embodiments, the unitary structures are formed from an additive manufacturing process. 
     Referring now to  FIG.  19   , an exemplary embodiment is provided depicting an operation of the engine  10 . The embodiment provided in  FIG.  19    is configured substantially similarly to the embodiment depicted and described with regard to  FIG.  16   . Operation of the system provided here may be based substantially as described with regard to embodiments of the engine  10  as depicted and described with regard to  FIGS.  1 - 6    and  FIGS.  7 A- 7 B . In  FIG.  19   , the flow of air  91  is received at the second location  272 , such as an opening provided through the outer manifold  2316 . The flow of air  91  is received into the conduit  1324  formed between the outer manifold  2316  and the inner manifold  1316 . The flow of air  91  is routed into the plenum  383  via the inlet opening  1381  formed through the inner manifold  1316 . The flow of air  91  is routed across the plurality of pins  334  and through the flowpath  330  (see  FIG.  17   ) into the inner wall conduit  1326  (see  FIG.  17   ). 
     In one embodiment, such as depicted in  FIG.  19   , the flow of air  92  may egress from the inner wall conduit  1326  to outside of the casing  300  or engine  10 , such as depicted via arrows  93  through opening  1380 . The flow of air  93  may egress heat or thermal energy from the thermal control rings  314  to an atmospheric condition, or to an under-casing or under-cowl area. 
     Referring now to  FIG.  20   , a perspective view of a portion of the engine  10  is provided. The embodiment provided in  FIG.  20    is configured substantially similarly to the embodiment described with regard to  FIGS.  16 - 19   . In particular,  FIG.  20    depicts a plurality of discrete flowpaths  330  extended in adjacent circumferential arrangement through the thermal control rings  314 . A plurality of outlet openings  3182  is formed through the inner portion  347  of the inner manifold  1316  corresponding to the plurality of flowpaths  330  and outlet openings  338  at the thermal control rings  314 . The engine  10  may accordingly form a plurality of flowpaths  330  and outlet openings  338  at the thermal control rings  314  in adjacent arrangement along the circumferential direction C corresponding to the plurality of outlet openings  3182  formed through the inner portion  347  of the inner manifold  1316 . Such arrangement may allow for the flow of air  92  to egress from within the thermal control ring  314  into the inner wall conduit  1326 . 
     Referring now to  FIG.  21   , a side cross-sectional view of the embodiment provided in  FIG.  20    is provided. The embodiment in  FIG.  21    further depicts the inner wall conduit  1326  in fluid communication with the second cavity  1322  positioned at the turbine frame  308 . An opening  3112  is formed through the turbine frame  308  to allow the flow of air  92  to egress into thermal communication with the turbine frame  308 . 
     Referring briefly now back to  FIG.  12    and  FIGS.  13 A- 13 D , additional aspects of the present disclosure are described.  FIG.  12    provides a partial circumferential view of an embodiment of the manifold assembly  316 .  FIGS.  13 A- 13 D  furthermore provide sectional views of the embodiment depicted in  FIG.  12    (labels for each of  FIGS.  13 A- 13 D  are indicated in  FIG.  12   ). As previously described, various embodiments of the manifold assembly  316  are formed via one or more additive manufacturing processes. Referring particularly to  FIG.  12    and the close-up view of  FIG.  13 C , in various embodiments, a member  3316  is extended to the inner manifold  1316  and the outer manifold  2316 . The member  3316  is extended at an acute angle (e.g., a V-, Z-, or other angled cross-section) from the inner manifold  1316  to the outer manifold  2316 . In various embodiments, the member  3316  is extended along a first direction, depicted schematically via arrows  95 , and a second direction opposite of the first direction, depicted schematically via arrows  96 . 
     Embodiments of the improved turbine casing  300 , turbine section  27 , and engine  10  provided herein allow for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. Embodiments of the engine  10 , the casing  300 , and manifold assembly  316  provided herein include integral, unitary structures, such as the casing extended over the stages of the high speed turbine, or further including the inter-turbine frame, or further including all or part of the manifold, such as may be formed by additive manufacturing processes that would not have heretofore been possible or practicable. Embodiments depicted and described herein allow for improved and advantageous positioning of thermal control rings  314 , flowpaths  330  therethrough, and the plurality of pins  334 , for improved clearance control response, improved formation and positioning of openings, passages, and conduits to allow for more efficient heat transfer fluid utilization and movement, and reduced weight, such as via obviating flanges and sub-assemblies into integral components. Particular combinations of these features allow for improved heat transfer properties and reduced thermal gradients. Improved heat transfer properties particularly include lowering a heat transfer coefficient at certain features, such as the plurality of walls, body, pins, and/or flowpaths forming the thermal control rings  314 , in contrast to known clearance control systems. Such improvements may mitigate or eliminate undesired or excessive deformation, ovalization, bowing, or other changes in geometry of the casing  300  that may adversely affect deflections or result in undesired contact to the turbine rotor blades  58  at the high speed turbine  28 . 
     Embodiments of the engine  10  and the casing  300  provided herein include an integral, unitary casing for the high speed turbine  28  together with a turbine center frame or mid-turbine frame  308 , formed by the outer casing wall  312  and the plurality of vanes  310  and positioned downstream along the core gas flowpath  70  of the high speed turbine  28  and upstream along the core gas flowpath  70  of a low- or intermediate-pressure turbine, such as depicted at turbine  30 . Embodiments provided herein further include e.g., an integral, unitary clearance control manifold configured to provide heat transfer fluid to thermal control rings. The integral, unitary structures may further allow for improved positioning of the thermal control rings relative to the turbine rotors, such as to provide improved clearance control across the turbine rotor assembly. 
     It should be appreciated that the conduits  110 ,  120 ,  123 , flow control devices  130 , or heat exchangers  141 ,  142  depicted and described with regard to  FIGS.  1 - 6    may be provided to the casing  300 , manifold assembly  316 , and other structures depicted and described with regard to  FIGS.  8 - 21   . However, various embodiments of the engine  10  provided herein may include one or more of the conduits  110 ,  120 ,  123 , flow control devices  130 , or heat exchangers  141 ,  142  providing flows of air to any appropriate clearance control system, turbine section, or bearing assembly. Such structures, when combined with any appropriate clearance control system, turbine section, or bearing assembly, may provide one or more advantages and benefits described herein. Alternatively, various embodiments of the engine  10  provided herein may include one or more of the casings  300  or the manifold assemblies  316  receiving flows of air from any appropriate conduits, passageways, flowpaths, tubes, or other structures. Such structures, when combined with any appropriate conduit or heat exchanger, may provide one or more advantages and benefits described herein. Benefits and advantages described with regard to either the conduits, flow control devices, heat exchangers, casings, or manifolds, when combined together, may compound such benefits and advantages described herein. 
     Embodiments of the conduits  110 ,  120 ,  123  and heat exchangers  141 ,  142  provided herein may be formed, at least in part, by one or more additive manufacturing processes such as described herein. For instance, the first heat exchanger  141  may be integrally formed with the first conduit  110 , or the second heat exchanger  142  may be integrally formed with the second conduit  120 , or portions thereof. In another instance, all or part of the first conduit  110 , including one or more inlet manifolds  111 , outlet manifolds  112 , or collectors  115  may be integrally formed as a single, unitary component. In still another instance, all or part of the second conduit  120 , including one or more inlet portions  121  or outlet portions  122  may be formed as a single, unitary component. Still further, certain combinations of portions of the first conduit  110 , second conduit  120 , and third conduit  123  may be formed integrally to one another. For instance, the outlet manifold  112  may be formed as a single, unitary component with the inlet portion  121 . In another instance, casings surrounding the compressor section  21  may be formed integrally with the inlet manifold  111 . The collector  115  may be formed integrally with the first heat exchanger  141 . The second heat exchanger  142  may be formed integrally with the outlet portion  122 . 
     This written description uses examples to disclose the preferred embodiments, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure 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. 
     Further aspects of the disclosure are provided by the subject matter of the following clauses: 
     A gas turbine engine, wherein the gas turbine engine defines an axial direction, a centerline axis parallel to the axial direction, a radial direction extended from the centerline axis, and a circumferential direction relative to the centerline axis, the gas turbine engine comprising: a first turbine rotor assembly comprising a plurality of first turbine rotor blades extended within a gas flowpath; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises an outer casing wall extended around the first turbine rotor assembly; a plurality of vanes extended from the outer casing wall and within the gas flowpath at a location aft of the first turbine rotor assembly; and a thermal control ring positioned outward along the radial direction from the outer casing wall, and wherein the thermal control ring comprises a body and a plurality of pins, and wherein the plurality of pins extend between the outer casing wall and the body. 
     The engine of one or more of these clauses, wherein the plurality of pins define a space between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allows for a flow of fluid through the space. 
     The engine of one or more of these clauses, wherein a flowpath is formed through the body of the thermal control ring. 
     The engine of one or more of these clauses, wherein a flowpath is extended along the radial direction through the body, and wherein the flowpath provides fluid communication between the space formed by the plurality of pins and an outlet opening positioned opposite the body from the space. 
     The engine of one or more of these clauses, wherein the flowpath is a plurality of discrete flowpaths in adjacent arrangement along the circumferential direction. 
     The engine of one or more of these clauses, the engine comprising: an inner manifold forming a double wall structure comprising an outer portion separated from an inner portion, wherein an inner wall conduit is formed between the outer portion and the inner portion, and wherein an outlet opening is formed through the inner portion corresponding to the outlet opening formed through the thermal control ring to allow for a flow of air from the flowpath at the thermal control ring into the inner wall conduit. 
     The engine of one or more of these clauses, wherein the inner manifold surrounds the thermal control ring along the circumferential direction and the axial direction. 
     The engine of one or more of these clauses, wherein the inner manifold is connected to the outer casing wall forward of the plurality of vanes. 
     The engine of one or more of these clauses, wherein the inner manifold forms an inlet opening through the double wall structure. 
     The engine of one or more of these clauses, wherein the inner manifold comprises an inlet opening wall forming an inlet opening flowpath fluidly separate from the inner wall conduit. 
     The engine of one or more of these clauses, wherein a seal is positioned in contact with the inner portion of the inner manifold and the outer surface of the thermal control ring. 
     The engine of one or more of these clauses, wherein a plenum wall is extended from the inner manifold and surrounding the thermal control ring. 
     The engine of one or more of these clauses, wherein a cavity is formed between an outer surface of the thermal control ring and the plenum wall. 
     The engine of one or more of these clauses, wherein a seal is positioned in contact with the outer surface of the thermal control ring and the plenum wall. 
     The engine of one or more of these clauses, wherein the inner portion of the inner manifold forms the outlet openings in adjacent circumferential arrangement radially outward of corresponding outlet openings of the thermal control ring. 
     The engine of one or more of these clauses, wherein the inner manifold is extended forward along the axial direction of the plurality of vanes, and wherein the inner manifold is connected to the outer casing wall forward of the plurality of vanes. 
     The engine of one or more of these clauses, wherein the outer casing wall, the plurality of pins, and the body of the thermal control rings are a unitary, integral structure. 
     A casing for a gas turbine engine defining an axial direction, a radial direction, a circumferential direction, and a gas flowpath, the gas turbine engine including a first turbine rotor assembly comprising a plurality of first turbine rotor blades extended within the gas flowpath, the casing comprising an outer casing wall configured to extend around the first turbine rotor assembly when the casing installed in the gas turbine engine; a plurality of vanes extending from the outer casing wall and configured to extend into the gas flowpath at a location aft of the first turbine rotor assembly when the casing is installed in the gas turbine engine; and a thermal control ring positioned outward along the radial direction from the outer casing wall, and wherein the thermal control ring comprises a body and a plurality of pins, and wherein the plurality of pins extend between the outer casing wall and the body. 
     The casing of one or more of these clauses, wherein the plurality of pins define a space between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allows for a flow of fluid through the space. 
     The casing of one or more of these clauses, wherein a flowpath is formed through the body of the thermal control ring, and wherein the flowpath provides fluid communication between the space formed by the plurality of pins and an outlet opening positioned opposite the body from the space.