Patent Publication Number: US-2023147089-A1

Title: Clearance control structure for a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to Polish Patent Application No. P. 439447, filed Nov. 5, 2021, which is a non-provisional application, and wherein the above application is hereby incorporated by reference in its entirety. 
     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. 
     FIELD 
     The present subject matter relates generally to clearance control structures for gas turbine engines. The present subject matter relates particularly to clearance control structures for turbine sections of gas turbine engines. 
     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. Conventional 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 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     An aspect of the disclosure is directed to a gas turbine engine having a first turbine rotor assembly including a plurality of first turbine rotor blades extended within a gas flowpath. A second turbine rotor assembly is positioned aft along the gas flowpath of the first turbine rotor assembly. The second turbine rotor assembly is rotatably separate from the first turbine rotor assembly. A casing surrounds the first turbine rotor assembly. The casing includes an outer casing wall extended forward of the first turbine rotor assembly and aft of the first turbine rotor assembly. The casing includes a plurality of vanes extended from the outer casing wall and through the gas flowpath aft of the first turbine rotor assembly and forward of the second turbine rotor assembly. The casing includes a plurality of walls forming thermal control rings extended outward along the radial direction from the outer casing wall. The outer casing wall and the thermal control rings is a unitary, integral structure. 
     Another aspect of the present disclosure is directed to a gas turbine engine having the first turbine rotor assembly, the second turbine rotor assembly, and the casing. An inner manifold wall surrounds the plurality of walls at the casing along the circumferential direction and the axial direction. The inner manifold wall is extended forward along the axial direction of the plurality of vanes, and the inner manifold wall is connected to the outer casing wall forward of the plurality of vanes. An outer manifold wall surrounds the inner manifold wall. In certain embodiments, the inner manifold wall and the outer manifold wall together form a unitary, integral structure. 
     These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG.  1    is an exemplary schematic cross sectional view of a turbomachine engine including a turbine section and casing in accordance with aspects of the present disclosure; 
         FIGS.  2 - 5    are exemplary schematic cross sectional view of embodiments of a portion of a turbine section and casing in accordance with aspects of the present disclosure; 
         FIG.  6    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.  6 A- 6 D  are exemplary sectional views of an embodiment of the manifold provided in  FIG.  6   ; 
         FIG.  7    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; and 
         FIG.  8    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. 
     
    
    
     Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention. 
     DETAILED DESCRIPTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. 
     The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     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, 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”). 
     An improved turbine casing is provided 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 pressure turbine casing and turbine center frame or mid-turbine frame positioned downstream of the high pressure 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 integra, 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 propulsion systems, auxiliary power units (APU), industrial gas turbines, hybrid-electric gas turbines, 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. In general, the engine  10  may include a core engine  14  disposed downstream from a fan section  16 . 
     The core engine  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 relationship, a compressor section having a booster or low speed compressor  22 , a high speed compressor  24 , a combustion section  26 , a turbine section including a high speed turbine  28 , a low speed turbine  30  (e.g., including vanes  116  and rotor blades  118 ), and a jet exhaust nozzle section  32 . 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 spool  36  may also be connected to a fan spool or shaft  38  of the fan section  16 . In particular embodiments, the low spool  36  may be connected directly to the fan spool  38  such as in a direct-drive configuration. In alternative configurations, the low spool  36  may be connected to the fan spool  38  via a speed reduction device  37  such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within engine  10  as desired or required. 
     It should be appreciated that the terms “low” and “high”, or their respective comparative degrees (e.g., −er, where applicable), when used with compressor, turbine, shaft, or spool components, 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. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low turbine” or “low speed turbine” may refer to the lowest maximum rotational speed turbine within a turbine section, a “low compressor” or “low speed compressor” may refer to the lowest maximum rotational speed compressor within a compressor section, a “high turbine” or “high speed turbine” may refer to the highest maximum rotational speed turbine within the turbine section, and a “high compressor” or “high speed compressor” may refer to the highest maximum rotational speed compressor within the compressor section. Similarly, the low speed spool refers to a lower maximum rotational speed than the high speed spool. It should further be appreciated that the terms “low” or “high” in such aforementioned regards may additionally, or alternatively, be understood as relative to minimum allowable speeds, or minimum or maximum allowable speeds relative to normal, desired, steady state, etc. operation of the engine. 
     Although depicted and described as a two-spool engine including the high speed spool separately rotatable from the low speed spool, it should be appreciated that the engine  10  may be configured as a three-spool engine including the high speed spool, the low speed spool, and an intermediate speed spool positioned in serial flow arrangement between the high speed spool and the low speed spool. It should further be appreciated that the low speed turbine or second turbine rotor assembly described herein generally refers to a separately rotatable spool downstream of the high speed turbine or first turbine rotor assembly. As such, the second turbine rotor 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 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 core engine  14 . It should be appreciated by those of ordinary skill in the art that the nacelle  42  may be configured to be supported relative to the core engine  14  by a plurality of circumferentially-spaced outlet guide vanes  44 . Moreover, a downstream section  46  of the nacelle  42  (downstream of the guide vanes  44 ) may extend over an outer portion of the core engine  14  so as to define a bypass airflow passage  48  therebetween. It should further be appreciated by those of ordinary skill in the art that certain embodiments of the engine may omit the nacelle  42 , such as to form a propfan or open rotor engine. Additionally, it should be appreciated by those of ordinary skill in the art that embodiments of the core engine  14  provided herein may be applied to other gas turbine engine configurations such as provided herein. 
       FIG.  2    provides an enlarged cross sectioned view of the turbine section portion of the core engine  14  as shown in  FIG.  1   , as may incorporate various embodiments of the present disclosure. As shown in  FIG.  2   , a first turbine rotor assembly is formed by the high speed turbine  28 . The first turbine rotor assembly includes a plurality of first turbine rotor blades  58  extended within the core gas flowpath  70 . A first stage  50  of the first turbine rotor assembly 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 HP spool  34  ( FIG.  1   ). The stator vanes  54 ,  64  and the turbine rotor blades  58 ,  68  at least partially define a core gas flowpath  70  for routing combustion gases from the combustion section  26  ( FIG.  1   ) through the high speed turbine  28 . 
     As further shown in  FIG.  2   , the high speed turbine 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 an active clearance control (“ACC”) system (not labeled). The ACC is used to minimize radial blade tip clearance CL between the outer blade tip and the shroud, particularly during cruise operation of the engine. 
     Downstream along the core gas flowpath  70 , or aft of the high speed turbine  28 , is a second turbine rotor assembly formed by the low speed turbine  30 . As previously described herein, the second turbine rotor assembly is rotatably separate from the first turbine rotor assembly, such as described in regard to the high speed turbine  28  and the low speed turbine  30  above with reference to  FIG.  1   . 
     A 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 rotor assembly formed by the high speed turbine  28  and forward of the second turbine rotor 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  ( FIG.  1   ). The outer casing wall  312  is extended along an 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 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.  2   . 
     It should be appreciated that conventional turbine casings include separable or joined flanges, such as bolted flanges or welded flanges, between a high pressure turbine casing and a downstream casing, such as an inter-turbine frame, mid-turbine frame, intermediate pressure turbine casing, or low pressure 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  FIGS.  3 - 4   , the forward thermal control rings  3141  are positioned in alignment along the axial direction A to the first stage of rotor blades  58 . 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  FIGS.  3 - 4   , 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 . 
     The forward and aft thermal control rings  3141  and  3142  are provided to more effectively control blade tip clearance CL (shown in  FIG.  2   ) 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 provides 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 , 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 overcome issues associated with integrating structures while providing unexpected benefits. In one instance, axially-extended casings are generally 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  316  surrounds the thermal control rings  314  along the circumferential direction C and the axial direction A. The manifold  316  is configured to provide a flow of fluid, such as relatively cool air from the compressor section, to the thermal control rings  314 . 
     Referring now to the  FIGS.  2 - 5   , and  FIG.  7   , further exemplary embodiments are provided. The embodiment depicted in  FIG.  2   ,  FIG.  3   , and  FIG.  7    may be configured similarly as one another, such as further described below.  FIGS.  4 - 5    provide views of flows of fluid and openings at various cross-sections of  FIG.  3   . Each of the embodiments may be formed via one or more manufacturing methods known in the art. In  FIG.  7   , 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.  2 - 5    and  FIG.  7   , the manifold  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.  7   , the manifold  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.  2   , the manifold  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.  2   , the manifold  316  includes a plurality of concentric walls integrally formed and surrounding the outer casing wall  312 . In certain embodiments, the manifold  316  includes an inner manifold wall  1316  radially inward of and concentric to an outer manifold wall  2316 . In still certain embodiments, the inner manifold wall  1316  is a double wall structure concentric to the outer manifold wall  2316 . 
     Referring particularly to  FIGS.  3 - 5   , 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  316 . In a still particular embodiment, the corrugated feature  399  may be formed at an inner manifold wall  1316  or an outer manifold wall  2316 . The corrugated feature  399  may allow for the unitary, integral formation of the manifold  316  with the outer casing wall  312 , such as described in various embodiments herein. 
     Referring briefly to  FIG.  8   , and in conjunction with  FIGS.  2 - 7   , the manifold  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 fluid  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 fluid  91 , depicted schematically via flow of fluid  92 , to egress the first cavity  1321  and enter an inner wall conduit  1326  such as described further below. 
     Referring to  FIGS.  3 - 7   , in particular embodiments, the manifold  316  includes an inner manifold wall  1316  surrounding the thermal control rings  314  along the circumferential direction C and the axial direction A. The manifold  316  may further include the outer manifold wall  2316  surrounding the inner manifold wall  1316 , as discussed above. A passage wall  1318  is extended to the outer manifold wall  2316  and the inner manifold wall  1316  to form a passage  1320  within the passage wall  1318 . 
     In certain embodiments, such as depicted in  FIG.  2   , the outer manifold wall  2316  of the manifold  316  is extended along the axial direction A at or aft the plurality of vanes  310 . The outer manifold wall  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.  3 - 5   , the inner manifold wall  1316  is extended to a location forward along the axial direction A of the plurality of vanes  310 . The inner manifold wall  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 wall  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 is formed between the inner manifold wall  1316  and the outer casing wall  312 . The thermal control rings  314  are surrounded by the inner manifold wall  1316  at a location within the first cavity  1321  between the inner manifold wall  1316  and the outer casing wall  312 . The passage  1320  allows for fluid communication with the first cavity  1321  between the inner manifold wall  1316  and the outer casing wall  312 . The passage  1320  further allows for the flow of fluid  91  to enter into thermal communication with the thermal control rings  314 . 
     In various embodiments, a conduit  1324  is formed between the outer manifold wall  2316  and the inner manifold wall  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 wall  2316  to the inner manifold wall  1316  through the conduit  1324 . 
     Referring to  FIGS.  3 - 5   , and further in regard to  FIG.  7   , the conduit  1324  is extended in fluid communication through one or more of the plurality of vanes  310 .  FIG.  4    and  FIG.  7    particularly depict the flow of fluid  91  entering into thermal communication and fluid communication with the thermal control rings  314  in the first cavity  1321 .  FIG.  4    particularly depicts the flow of fluid  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.  5    and  FIG.  7    particularly depict the flow of fluid  92  egressing from the first cavity  1321  through the conduit  1324  and then in serial flow through one or more of the 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.  6 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 to  FIG.  7   , and further depicted in the detailed perspective view in  FIG.  8   , in certain embodiments, the inner manifold wall  1316  is a double wall structure forming an inner wall conduit  1326  between the double wall structure of the inner manifold wall  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 gas flowpath  70 . In such embodiments, the unitary, integral casing  300 , or furthermore integral to embodiments of the manifold  316 , allow for separate flows into the plurality of vanes  310 . Particularly, the flow of fluid  91  enters the conduit  1324  from a compressor section or other fluid source. A portion of the flow of fluid  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 fluid  92  then flows into one or more of the plurality of vanes  310 . Furthermore, another portion of the flow of fluid  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 manifold  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 gas flowpath  70  forms the outer shroud segment  77  of the shroud assembly  72 . The outer shroud segment  77  is exposed to the 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 gas flowpath  70 . 
     Referring now to  FIG.  6   , a partial circumferential view of an embodiment of the manifold  316  is provided.  FIGS.  6 A- 6 D  furthermore provide sectional views of the embodiment depicted in  FIG.  6   . As previously described, various embodiments of the manifold  316  are formed via one or more additive manufacturing processes. Referring particularly to the close-up view of  FIG.  6 C , in various embodiments, a member  3316  extended to the inner manifold wall  1316  and the outer manifold wall  2316 . The member  3316  is extended at an acute angle (e.g., a V-, Z-, or other angled cross-section) from the inner manifold wall  1316  to the outer manifold wall  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  and engine  10  provided herein allow for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. Embodiments of the engine  10 , casing  300 , and manifold  316  provided herein include integral, unitary structures, such as the casing extended over the stages of the high pressure 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  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 forms 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 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, formed by the outer casing wall  312  and vanes  310  and positioned downstream along the gas flowpath  70  of the high speed turbine  28  and upstream along the 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. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims. 
     Further aspects of the invention are provided by the subject matter of the following clauses: 
     1. A gas turbine engine, wherein the engine defines an axial direction co-directional to a centerline axis, a radial direction extended from the centerline axis, and a circumferential direction relative to the centerline axis, the engine comprising a first turbine rotor assembly comprising a plurality of first turbine rotor blades extended within a gas flowpath; a second turbine rotor assembly positioned aft along the gas flowpath of the first turbine rotor assembly, wherein the second turbine rotor assembly is rotatably separate from the first turbine rotor assembly; a casing surrounding the first turbine rotor assembly, wherein the casing comprises an outer casing wall extended forward of the first turbine rotor assembly and aft of the first turbine rotor assembly, and wherein the casing comprises a plurality of vanes extended from the outer casing wall and through the gas flowpath aft of the first turbine rotor assembly and forward of the second turbine rotor assembly, and further wherein the casing comprises a plurality of walls forming thermal control rings extended outward along the radial direction from the outer casing wall, and wherein the outer casing wall and the thermal control rings is a unitary, integral structure. 
     2. The gas turbine engine of any one or more clauses herein, wherein the plurality of walls comprises a plurality of axially-spaced stages, wherein the plurality of axially-spaced stages corresponds to an axial position of each respective first turbine rotor stage. 
     3. The gas turbine engine of any one or more clauses herein, the engine comprising a manifold surrounding the plurality of walls along the circumferential direction and the axial direction, wherein the manifold is configured to provide a flow of fluid to the plurality of walls, and wherein the manifold is a unitary, integral structure. 
     4. The gas turbine engine of any one or more clauses herein, wherein the manifold is extended along the axial direction forward and aft of the plurality of axially-spaced stages of the plurality of walls. 
     5. The gas turbine engine of any one or more clauses herein, wherein the manifold comprises a plurality of openings surrounding the plurality of walls at the casing, 
     6. The gas turbine engine of any one or more clauses herein, wherein the manifold comprises an inner manifold wall surrounding the plurality of walls at the casing along the circumferential direction and the axial direction, and an outer manifold wall surrounding the inner manifold wall, wherein a passage wall is extended to the outer manifold wall and the inner manifold wall to form a passage therewithin, and wherein the passage is in fluid communication with a first cavity between the inner manifold wall and the outer casing wall. 
     7. The gas turbine engine of any one or more clauses herein, wherein a conduit is formed between the outer manifold wall and the inner manifold wall, wherein the conduit is in fluid communication with the first cavity, and wherein the passage wall separates the conduit from the passage within the passage wall. 
     8. The gas turbine engine of any one or more clauses herein, wherein the conduit is extended in fluid communication to a second cavity formed between the outer casing wall and the shroud assembly. 
     9. The gas turbine engine of any one or more clauses herein, wherein the conduit is extended in fluid communication through one or more of the plurality of vanes. 
     10. The gas turbine engine of any one or more clauses herein, wherein the manifold, the outer casing wall, and the plurality of walls of the casing is a unitary, integral structure. 
     11. The gas turbine engine of any one or more clauses herein, wherein the outer manifold wall is extended aft along the axial direction of the plurality of vanes, and wherein the outer manifold wall is connected to the outer casing wall aft of the plurality of vanes. 
     12. The gas turbine engine of any one or more clauses herein, wherein the inner manifold wall is extended forward along the axial direction of the plurality of vanes, and wherein the inner manifold wall is connected to the outer casing wall forward of the plurality of vanes. 
     13. The gas turbine engine of any one or more clauses herein, wherein the inner manifold wall comprises a double wall structure forming an inner wall conduit. 
     14. The gas turbine engine of any one or more clauses herein, wherein the inner wall conduit is extended in fluid communication to a second cavity formed between the outer casing wall and an outer wall of the gas flowpath. 
     15. The gas turbine engine of any one or more clauses herein, wherein the manifold is extended along the axial direction forward and aft of the plurality of axially-spaced stages of the plurality of walls, and wherein the manifold is extended aft along the axial direction of the plurality of vanes. 
     16. The gas turbine engine of any one or more clauses herein, wherein the manifold comprises an inner manifold wall surrounding the plurality of walls at the casing along the circumferential direction and the axial direction; an outer manifold wall surrounding the inner manifold wall, wherein a passage wall is extended to the outer manifold wall and the inner manifold wall to form a passage therewithin, and wherein the passage is in fluid communication with a first cavity between the inner manifold wall and the outer casing wall; and a member extended to the inner manifold wall and the outer manifold wall, wherein the member is extended at an acute angle from the inner manifold wall to the outer manifold wall along a first direction and a second direction opposite of the first direction. 
     17. The gas turbine engine of any one or more clauses herein, wherein the casing forms a corrugated feature extended along the axial direction. 
     18. A gas turbine engine, wherein the engine defines an axial direction co-directional to a centerline axis, a radial direction extended from the centerline axis, and a circumferential direction relative to the centerline axis, the engine comprising a first turbine rotor assembly comprising a plurality of first turbine rotor blades extended within a gas flowpath; a second turbine rotor assembly positioned aft along the gas flowpath of the first turbine rotor assembly, wherein the second turbine rotor assembly is rotatably separate from the first turbine rotor assembly; a casing surrounding the first turbine rotor assembly, wherein the casing comprises a unitary, integral outer casing wall extended forward of the first turbine rotor assembly and aft of the first turbine rotor assembly, and wherein the casing comprises a plurality of vanes extended from the outer casing wall and through the gas flowpath aft of the first turbine rotor assembly and forward of the second turbine rotor assembly, and further wherein the casing comprises a plurality of walls extended outward along the radial direction from the outer casing wall, and wherein the outer casing wall and the plurality of walls is a unitary, integral structure; an inner manifold wall surrounding the plurality of walls at the casing along the circumferential direction and the axial direction, wherein the inner manifold wall is extended forward along the axial direction of the plurality of vanes, and wherein the inner manifold wall is connected to the outer casing wall forward of the plurality of vanes; and an outer manifold wall surrounding the inner manifold wall, wherein the outer manifold wall and the inner manifold wall together form a unitary, integral structure. 
     19. The gas turbine engine of any one or more clauses herein, wherein a passage wall is extended to the outer manifold wall and the inner manifold wall to form a passage therewithin, and wherein the passage is in fluid communication with a first cavity between the inner manifold wall and the outer casing wall. 
     20. The gas turbine engine of any one or more clauses herein, wherein the inner manifold wall forms a double wall structure, and wherein an inner wall conduit is formed within the double wall structure in fluid communication with a first cavity formed between the inner manifold wall and the outer casing wall.