Gas turbine engine with clearance control system

Abstract

A gas turbine engine comprising: a first turbine rotor assembly comprising a plurality of first turbine rotor blades extending within a gas flow path; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises an outer casing wall extending around the first turbine rotor assembly, a plurality of vanes extending from the outer casing wall and extending within the gas flow path at a location aft of the first turbine rotor assembly, and a thermal control ring positioned radially outward of 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.

Description

[0001] Cross-references to related applications

[0002] This application claims priority to Polish patent application number P.439449, filed on November 5, 2021, as a non-provisional application, the entire contents of which are incorporated herein by reference.

[0003] Government-funded research

[0004] The project that led to this application has been funded by the EU Clean Sky 2 Research and Innovation Programme under grant agreement CS2-ENG-GAM-2014-2015-01. Technical Field

[0005] This topic specifically relates to gas turbine engines, including gap control structures. Background Technology

[0006] Housings for gas turbine engines, such as turbine section housings surrounding a turbine section rotor, generally require separable flanges and assembled housing and manifold portions due to internally and externally mounted components. These components generally include brackets or hangers for the turbine shield, or flanges for multiple housings. Additionally, because the turbine housing surrounds the turbine rotor, excessive deformation, thermal expansion or contraction, or bending can lead to excessive friction and undesirable contact with the turbine rotor, which may result in performance or operability losses. Some housings may include components that limit deformation or displacement during engine operation and thermal cycling via separable flanges. However, the inventors of this disclosure have found that such designs require components and parts that increase engine weight. Furthermore, the inventors of this disclosure have found that such designs can further suppress the inclusion or placement of thermal control structures for more effective clearance control.

[0007] Thus, the inventors of this disclosure have found a need to overcome these limitations and to provide turbine housings that improve thermal control, improve engine efficiency, and reduce weight. Summary of the Invention

[0008] According to one aspect of the present disclosure, there is provided a gas turbine engine, wherein the gas turbine engine defines an axial direction, a centerline axis parallel to the axial direction, a radial direction extending 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 extending within a gas flowpath; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises: an outer casing wall extending around the first turbine rotor assembly; a plurality of vanes extending from the outer casing wall and extending within the gas flowpath at a location aft of the first turbine rotor assembly; and a thermal control ring positioned radially outward of 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; wherein the plurality of pins define a gap between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid flow through the gap, and a flowpath is formed through an interior of the body of the thermal control ring.

[0009] According to another aspect of the present disclosure, there is provided a casing for a gas turbine engine, the gas turbine engine defining an axial direction, a radial direction, a circumferential direction, and a gas flowpath, the gas turbine engine comprising a first turbine rotor assembly comprising a plurality of first turbine rotor blades extending within the gas flowpath, the casing comprising: an outer casing wall configured to extend around the first turbine rotor assembly when the casing is 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 radially outward of 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; wherein the plurality of pins define a gap between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid flow through the gap, and a flowpath is formed through an interior of the body of the thermal control ring. BRIEF DESCRIPTION OF DRAWINGS

[0010] In the description of the specification, which follows, a complete and enabling disclosure is set forth, including best modes for practicing the application, as presently contemplated, which includes specific embodiments and examples of the disclosure, to one of ordinary skill in the art, and it will be apparent to one of ordinary skill in the art that various changes and modifications can be made thereto without departing from the intended spirit and scope of the disclosure.

[0011] Figure 1is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of the present disclosure;

[0012] Figure 2 is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of the present disclosure;

[0013] Figure 3 is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of the present disclosure;

[0014] Figure 4 is a schematic cross-sectional view of a portion of an embodiment of a gas turbine engine according to aspects of the present disclosure;

[0015] Figure 5 is a schematic cross-sectional view of a portion of an embodiment of a gas turbine engine according to additional aspects of the present disclosure;

[0016] Figure 6 is a perspective view of a portion of an embodiment of a gas turbine engine according to aspects of the present disclosure;

[0017] Figures 7A-7B depicts a flowchart outlining steps of a method for operating an engine according to aspects of the present disclosure;

[0018] Figures 8-11 is an exemplary schematic cross-sectional view of an embodiment of a portion of a turbine section and casing according to aspects of the present disclosure;

[0019] Figure 12 is an exemplary perspective view of an embodiment of a portion of a manifold of a turbine section according to aspects of the present disclosure;

[0020] Figures 13A-13D is Figure 12 an exemplary cross-sectional view of an embodiment of a manifold provided in

[0021] Figure 14 is an exemplary schematic cross-sectional view of an embodiment of a portion of a turbine section and casing according to aspects of the present disclosure;

[0022] Figure 15 is an exemplary perspective view of an embodiment of a portion of a manifold of a turbine section according to aspects of the present disclosure;

[0023] Figure 16 is an exemplary schematic cross-sectional view of an embodiment of a portion of a turbine section and casing according to aspects of the present disclosure;

[0024] Figure 17 is an exemplary schematic cross-sectional view of an embodiment of a portion of a turbine section and casing according to aspects of the present disclosure; Figure 16 is a detailed view of an exemplary schematic cross-sectional view of an embodiment of

[0025] Figure 18 is a top-down view of an exemplary embodiment of a plurality of pins of a thermal control ring according to aspects of the present disclosure;

[0026] Figure 19 is an exemplary schematic illustration of air flow through a turbine section and a casing according to aspects of the present disclosure; Figure 16

[0027] Figure 20 is a perspective view of a portion of an engine according to aspects of the present disclosure; and

[0028] Figure 21 is a cross-sectional view of an embodiment of an engine provided in Figure 20

[0029] Reference characters that are repeated throughout the specification and drawings are intended to designate the same or similar features or elements of the present disclosure. DETAILED DESCRIPTION

[0030] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations

[0031] 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. Furthermore, unless otherwise indicated, the description of all embodiments herein is to be taken in their generic sense and can be applied to a variety of generic embodiments.

[0032] As used herein, the terms "first," "second," and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of individual components.

[0033] 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.

[0034] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0035] The terms "coupled," "fixed," "attached" and the like, mean either directly coupled, fixed, or attached as applicable, or indirectly coupled, fixed or attached through one or more intermediate

[0036] ​​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 of the basic function to which it is directed. Accordingly, a value modified by a term or terms, such as "about", "approximately", and "substantially", are not limited to the precise value specified. In at least some instances, the approximating language can correspond to the precision of an instrument for measuring the value or the precision of the manner in which the value is digitized or otherwise captured. For example, the approximating language can refer to the precision to which digital measurements are conformed, or the precision to which differing characteristics are distinguishable. In some embodiments, the approximating language can refer to the precision of a measurement instrument or the precision of a method or machine for constructing or manufacturing a component and / or system.

[0037] Throughout this document and in the 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.

[0038] Pressure values and ranges are expressed in absolute pressure measurements (psia) or equivalent. Values and ranges of pressure provided herein can be converted to ranges of gauge pressure, or other units of pressure, or other units, measurements, or combinations thereof corresponding to the values and / or ranges disclosed herein.

[0039] The term "gross power output" refers to the maximum rated power output of an engine.

[0040] The term "operational envelope" refers to a cycle, mission, or set of maneuvers in which an engine can operate normally. In one embodiment, a landing-takeoff (LTO) cycle can define an operational envelope. An LTO cycle including one or more combinations of start-up, idle, takeoff, cruise, and approach engine operating conditions can collectively define an operational envelope. In various embodiments, the cruise condition defines a majority of the operational envelope, such as to define a majority of the operational time or duration of the engine operation. In certain embodiments, the cruise condition is between approximately 55% and 75% of the operational envelope. In other words, the cruise condition can define approximately 55% to approximately 75% of the engine operating duration from start-up to shut-down after an approach operating condition. In another embodiment, the cruise condition can define approximately 60% to approximately 70% of the engine operating duration.

[0041] The term "cruise operating condition" can further refer to a medium power engine operating condition. The term "takeoff operating condition" can refer to a full power condition and "idle operating condition" can 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 the gross power output of the engine. In still certain embodiments, the cruise condition corresponds to approximately 80% to 88% of the gross power output of the engine.

[0042] As used herein, "third stream" means a non-primary air stream capable of adding fluid energy to produce a minority of total propulsive system thrust. The pressure ratio of the third stream can be higher than the pressure ratio of a primary propulsive stream (e.g., a bypass or propeller driven propulsive stream). Thrust can be produced by a dedicated nozzle, or by mixing the airflow through the third stream with a primary propulsive stream or core air stream, for example, into a common nozzle.

[0043] In certain example embodiments, the operating temperature of the airflow through the third stream can be less than the maximum compressor discharge temperature of the engine, and more specifically, can 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 ambient temperature). In certain example embodiments, these operating temperatures can facilitate the transfer of heat to or out of the airflow through the third stream and separate fluid stream. Further, in certain example embodiments, the airflow through the third stream can contribute less than 50% of total engine thrust (and at least, for example, 2% of total engine thrust) in takeoff conditions, or more specifically, when operating in sea level rated takeoff power, static flight speed, 86 degrees Fahrenheit ambient temperature operating conditions.

[0044] Further, in certain example embodiments, aspects of the airflow through the third stream (e.g., airflow, mixing, or exhaust properties), and the aforementioned example percentages of contribution to total thrust thereby, can be adjusted passively during engine operation, or modified purposefully through the use of engine control features, such as fuel flow, electric machine power, variable stator, variable inlet guide vane, valve, variable exhaust port geometry, or fluidic features, to adjust or optimize overall system performance in a wide range of potential operating conditions.

[0045] The term "turbomachine" or "turbomachinery" refers to a machine that includes one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

[0046] 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, and the like.

[0047] The term "combustion section" refers to any hot addition system for a turbomachine. For example, the term combustion section can refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable hot addition assembly. In certain example embodiments, a combustion section can include an annular combustor, a can combustor, a can-annular combustor, a trapped vortex combustor (TVC), or other suitable combustion system, or combinations thereof.

[0048] The terms "low" and "high," or their respective comparative forms (e.g., lower, higher, as applicable), when used in connection with compressor, turbine, shaft, or spool components, etc., each refer to relative speeds within an engine, unless otherwise specified. For example, "low turbine" or "low speed turbine" defines a component configured to operate at a lower rotational speed (such as a maximum allowable rotational speed) than a "high turbine" or "high speed turbine" at the location of the engine.

[0049] The term "at" as used herein to refer to a position of a first object relative to a second object (e.g., the first object is located or positioned at the second object) means that the first object is positioned entirely or partially within the second object, the first object contacts the second object, or the first object is positioned closest to the second object (relative to any other surrounding related components).

[0050] One or more components of the turbomachine engines described below can 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 processes can allow such components to be integrally formed as a single unitary component, or as any suitable number of sub-components. In particular, additive manufacturing processes can allow such components to be integrally formed and include various features that would not be possible to achieve using prior manufacturing methods. For example, the additive manufacturing methods described herein can allow for the manufacture of passages, conduits, cavities, openings, housings, manifolds, double walls, heat exchangers, or other components, or specific positioning and integration of these components with unique features, configurations, thicknesses, materials, densities, fluid pathways, headers, and mounting structures that can not be possible or practical to achieve using prior manufacturing methods. Some of these features are described herein.

[0051] For example, suitable additive manufacturing techniques in accordance with the present disclosure include fused deposition modeling (FDM), selective laser sintering (SLS), 3D printing such as by inkjet, laser jet, and binder jet, 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.

[0052] Suitable powder materials for manufacturing the structures provided herein as unitary single structures include metal alloys, polymeric, or ceramic powders. Exemplary metal powder materials are stainless steel alloys, cobalt-chromium alloys, aluminum alloys, titanium alloys, nickel-based superalloys, and cobalt-based superalloys. Additionally, suitable alloys can include those alloys that have been designed to have good oxidation resistance, known as “superalloys,” which have acceptable strength at elevated operating temperatures in gas turbine engines, for example, 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 247LC, C263, 718, X-850, ECY 768, 282, X45, PWA1483, and CMSX (e.g., CMSX-4) single crystal alloys. Manufactured objects of the present disclosure can be formed with one or more selected crystalline microstructures, such as directionally solidified (“DS”) or single crystal (“SX”).

[0053] Embodiments of gas turbine engines are provided that include improved clearance control systems. The engines reduce weight and tubes, manifolds, or conduits external to the outer core casing or fan casing by reducing or eliminating air extracted from the fan bypass passage for cooling at the turbine section. Embodiments provided herein allow for engines without a fan casing, such as open rotor engines or propeller fan engines, to have and operate with improved clearance control, cooling systems, or air systems for the turbine section and / or bearing assemblies. It should be appreciated that while these embodiments can apply to turbine fan engines that include a nacelle and fan casing, embodiments provided herein allow for engines without a nacelle, fan casing, or other structure surrounding the fan section to receive air for turbine section cooling, clearance control, or bearing assemblies.

[0054] The improved gas turbine engines provided herein can additionally or alternatively allow for the removal of low pressure and / or low temperature air from the compressor section for cooling or clearance control at the turbine section and bearing assemblies. Certain clearance control systems can generally utilize high energy air (i.e., high pressure and / or high temperature air) such as from a rear stage of a high pressure compressor and mix with one or more other air sources such as from other compressor stages or from a fan air stream. Such high energy air reduces engine efficiency such as by removing energy from the thermodynamic and combustion processes or by requiring greater heat load reduction before the air is suitable for cooling or clearance control at the turbine section. Still further, certain clearance control systems can not be suitable for additionally providing air to bearing assemblies for cooling, buffer air, or other uses at the bearing assemblies.

[0055] Another aspect of the present disclosure is directed to improved turbine casings that allow for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. Embodiments of engines, casings, and manifolds provided herein include unitary single structures such as can be formed by additive manufacturing processes that have not been possible or feasible to date. 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 exclusion of flanges and subassemblies as unitary components. The particular combination of these features allows for improved heat transfer properties and reduced thermal gradients. The improved heat transfer properties include, among other things, lower heat transfer coefficients at certain features such as at the multiple walls forming the thermal control rings as provided herein. Such improvements can mitigate or eliminate undesirable or excessive distortion, ovalization, bowing, or other changes in casing geometry that can adversely affect deflection or cause undesirable contact with the turbine rotor.

[0056] Embodiments provided herein include, for example, a unitary single high speed turbine casing and a turbine center frame or intermediate turbine frame positioned downstream of the high speed turbine and upstream of a low pressure or intermediate pressure turbine. Embodiments provided herein further include, for example, a unitary single clearance control manifold configured to provide heat transfer fluid to the thermal control ring. The unitary single structure can further allow for improved positioning of the thermal control ring relative to the turbine rotor such as to provide improved clearance control across the turbine rotor assembly.

[0057] As used herein, the term "unitary single" used to describe a structure refers to a structure that is integrally formed from a continuous material or group of materials without seams, connecting joints, or the like. The unitary single structures described herein can be formed by additive manufacturing to have the described structure or, alternatively, by a casting process or the like.

[0058] Referring now to the drawings, Figure 1is a schematic cross-sectional view of an exemplary gas turbine engine 10, referred to herein as "engine 10," that can incorporate various embodiments of the present disclosure. Particular embodiments of the engine 10 can be configured as a turbofan, turboprop, turboshaft, or propfan gas turbine engine, or one or more gas turbine engines of a hybrid electric gas turbine engine, or other gas turbine engine configurations.

[0059] As shown in Figure 1 the engine 10 has a longitudinal or axial centerline axis 12 extending therethrough parallel to an axial direction A for purposes of reference. Generally, the engine 10 can include a turbomachine 14 disposed downstream of a fan section 16.

[0060] The engine 10 includes a compressor section 21 in a serial flow arrangement with a turbine section 27. The turbomachine 14 can generally include a substantially tubular outer casing 18 defining an annular inlet 20. The outer casing 18 can be formed of multiple casings. The outer casing 18 encloses the compressor section 21, a combustion section 26, and the turbine section 27 in a serial flow arrangement. In particular embodiments, the compressor section 21 includes a booster or low speed compressor 22 and a high speed compressor 24. In yet further particular embodiments, 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). An injection 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 can also be connected to a fan shaft or spool 38 of the fan section 16. In particular embodiments, the low speed spool 36 can be directly connected to the fan spool 38, such as in a direct drive configuration. In alternative configurations, as depicted by the dashed lines in Figure 1 the low speed spool 36 can be connected to the fan spool 38 via a gear assembly 37, such as in an indirect drive or geared drive configuration that allows for a higher or lower rotational speed of the fan spool 38 compared to the low speed spool 36, configuring the engine 10. Such a gear assembly can be included between any suitable shafts / spools within the engine 10 as desired or needed.

[0061] Although depicted and described as a dual-spindle engine including a high-speed spindle 34 capable of rotating separately from the low-speed spindle 36, it should be understood that engine 10 can be configured as a three-spindle engine, comprising a high-speed spindle 34, a low-speed spindle 36, and a third or intermediate-speed spindle positioned in a series flow arrangement between the high-speed spindle 34 and the low-speed spindle 36. Therefore, compressor section 21 may include an intermediate-speed compressor capable of rotating separately from the high-speed compressor 24 and the low-speed compressor 22. Similarly, turbine section 27 may include a third turbine assembly or intermediate-speed turbine capable of rotating separately from the high-speed turbine 28 and the low-speed turbine 30. The intermediate-speed compressor and the intermediate-speed turbine may be coupled together to fluidly form an intermediate-speed spindle between the high-speed and low-speed spindles.

[0062] It should further be understood that, in some embodiments, the low-speed turbine 30 or the second turbine assembly described herein generally refers to a separably rotatable spool downstream of the high-speed turbine or the first turbine assembly. Thus, the second turbine assembly may include a medium-speed turbine or a low-speed turbine positioned downstream of the high-speed turbine.

[0063] like Figure 1 As shown, fan section 16 includes one or more axially spaced stages of a plurality of fan blades 40 coupled to and extending radially outward from fan spool 38. An annular fan housing or nacelle 42 circumferentially surrounds at least a portion of fan section 16 and / or turbine 14. It should be understood that, for the depicted embodiment, nacelle 42 is supported relative to turbine 14 by a plurality of circumferentially spaced outlet guide vanes 44.

[0064] A bypass airflow passage 48 is formed downstream of one or more stages of the plurality of fan blades 40 and surrounds the outer portion of the turbine 14. In certain embodiments, such as Figure 1 As depicted, the bypass airflow passage 48 is defined in the downstream section 46 of the nacelle 42 (downstream of the outlet guide vane 44) and between the nacelle 42 and the outer portion of the turbine 14.

[0065] However, in other embodiments, it should be understood that the low-speed compressor 22 may form one or more stages of the fan section 16, such as Figure 3As such, bypass airflow passages 48 can generally include downstream of one or more stages of fan blades 40 and / or low speed compressor 22 and bypass or surround at least a portion of high speed compressor 24 with any flow path of bypass air flow 177 therethrough that provides thrust. Accordingly, certain embodiments of engine 10 provided herein can 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 fan blades 40 and / or low speed compressor 22 and upstream of at least a portion of high speed compressor 24, one or more of which are configured as a“third stream.”

[0066] Engine 10 includes a computing system 1210 configured to operate. Computing system 1210 is communicatively coupled to turbomachinery 14 and / or starter motor (not depicted) to regulate, modulate, maintain, vary, or link any one or more control surfaces to generate one or more embodiments of air flow, heat transfer fluid flow, and / or liquid and / or gaseous fuel in accordance with aspects of the disclosure provided herein. Computing system 1210 can generally correspond to any suitable processor-based device, including one or more computing devices. Certain embodiments of computing system 1210 include a full authority digital engine controller (FADEC), a digital engine controller (DEC), or other suitable computing device configured to operate engine 10.

[0067] Computing system 1210 can include one or more processors 1212 and one or more associated memory devices 1214 configured to perform various computer-implemented functions, such as steps of the methods described herein. As used herein, the term“processor” refers not only to integrated circuits capable of executing a software program, but also includes, without limitation, digital signal processor (DSP) chips, reduced instruction set (RISC) processor cores, general purpose (CISC) processor cores, baseband processor, and other

[0068] The computing system 1210 can include control logic 1216 stored in the memory 1214. The control logic 1216 can 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 those outlined in the one or more steps of the method 1000 further provided below. In still various embodiments, the memory 1214 can store charts, tables, functions, lookups, schedules, etc. corresponding to flow rates or velocities, pressures, or temperatures associated with the air flow, heat transfer fluid flow, or fuel flow 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 logical and / or virtual separate threads on the processor.

[0069] The computing system 1210 can also include a communication interface module 1230. In various embodiments, the communication interface module 1230 can include associated electronic circuitry for transmitting and receiving data. As such, the communication interface module 1230 of the computing system 1210 can be used to receive data from one or more control surfaces, sensors, measuring devices, or instruments, or calculations or measurements corresponding to one or more portions of the engine 10 provided herein, and can perform one or more steps of the method 1000 provided herein. The computing system 1210 can also include a network interface for communicating with other components of the engine 10, for example. The network interface can include any suitable components for interfacing with one or more networks, including for example, transmitters, receivers, ports, controllers, antennas, and / or other suitable components.

[0070] It should be understood that the communication interface module 1230 can be any combination of suitable wired and / or wireless communication interfaces, and as such can be communicatively coupled to one or more components of the device via wired and / or wireless connections. As such, the computing system 1210 can obtain, determine, store, generate, transmit, or operate any one or more steps of the methods described herein via a distributed network. For example, the network can include a SATCOM network, an ACARS network, an ARINC network, a SITA network, an AVICOM network, a VHF network, an HF network, a Wi-Fi network, a WiMAX network, a gate link network, etc.

[0071] Referring now to Figure 2 , exemplary embodiments of the open rotor configuration of the engine 10 are provided. Figure 1 The embodiments of the engine 10 provided in Figure 2 are configured substantially similar to those provided in Figure 1 However, in Figure 2In particular embodiments, the open-rotor configuration of engine 10 does not have a fan case or nacelle 42 (depicted in Figure 1 Downstream of the plurality of fan blades 40, or particularly downstream of the outlet guide vane 44, and radially outward of the outer portion of turbomachine 14, bypass airflow passages 48 are formed.

[0072] Reference is now made to Figure 3 Exemplary embodiments of an open-rotor configuration are provided in accordance with Figure 2 Reference is now made to Figure 3 The embodiments provided in In particular embodiments, the open-rotor configuration of engine 10 does not have a fan case or nacelle 42 (depicted in

[0073] Downstream of the plurality of fan blades 40, or particularly downstream of the outlet guide vane 44, and radially outward of the outer portion of turbomachine 14, bypass airflow passages 48 are formed. Figure 3 While Figure 1 depicts a three-flow or adaptive-cycle engine 10 having an open-rotor configuration, it should be appreciated that the adaptive-cycle configuration can also include a nacelle about the fan section, such as depicted and described with respect to In this manner, it should be further appreciated that while certain advantages and benefits provided herein can provide benefits for turbofan engines having a nacelle, embodiments and arrangements of components provided herein can overcome problems or challenges unique to open-rotor configurations.

[0074] Reference is now made to Figures 4-5 Exemplary embodiments of an open-rotor configuration are provided in accordance with Figures 1-3 Reference is now made to Figures 4-5A walled conduit, manifold, tube, or other structure forming a flow path is depicted that is configured to extract or receive an air flow schematically depicted via arrow 91 from the compressor section 21 and provide the air flow 91 to the turbine section 27. The air flow 91 provided to the turbine section 27 can be used to cool blades, vanes, shrouds, or other portions of the turbine section 27. In certain embodiments, the turbine section 27 includes a turbine frame 308 positioned between a first turbine assembly or high speed turbine 28 and a second turbine assembly or low speed turbine 30 in a series flow arrangement. In still particular embodiments, the bearing assembly 200 is included at the turbine frame 308. Thus, the turbine frame 308 can provide a static mounting or support structure at which the bearing assembly 200 is positioned to support rotation of one or more spools (e.g., the low speed spool 36 or the high speed spool 34). The turbine frame 308 further includes any suitable number of conduits, manifolds, or passages 309, or other structures for allowing at least a portion of the air flow 91 (e.g., depicted further below as air flow 193) to the bearing assembly 200. The air flow to the bearing assembly 200 can provide cooling or cushion air at the bearing assembly 200, such as to dampen vibrations from the spools or create a desired bearing or rotor clearance. In other embodiments, the air flow 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 ejection exhaust nozzle section 32.

[0075] The engine 10 includes a first conduit 110 extending in fluid communication from the compressor section 21 to the turbine section 27. The first conduit 110 is configured to communicate the air flow 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 flow path 70. In particular embodiments, the first conduit 110 provides the air flow 91 from the compressor section 21 to the turbine section 27 while bypassing the combustion section 26.

[0076] The first heat exchanger 141 is positioned in thermal communication with the airflow 91 through the first conduit 110. The first heat exchanger 141 is configured to receive heat or thermal energy from the airflow 91 through the first conduit 110. Accordingly, the first heat exchanger 141 is configured to cool the airflow 91 through the first conduit 110 prior to the airflow 91 being provided to the turbine section 27. The first heat exchanger 141 is configured as any suitable heat exchanger for extracting heat or thermal energy from the airflow 91 and receiving the heat or thermal energy, or transferring the heat or thermal energy to a heat transfer fluid, schematically depicted via arrow 221. Particular embodiments of the engine 10 can 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 a suitable cooler air or inert gas. The fluid system 220 provides the heat transfer fluid 221 in thermal communication with the airflow 91 via the first heat exchanger 141. In Figure 9 In the particular embodiment depicted (discussed in greater 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 can be provided and used in any suitable manner, including but not limited to as a lubricant for a bearing system, an anti-icing fluid, a fuel, or an actuating fluid.

[0077] Still referring to Figures 4-5 , the engine 10 includes a second conduit 120 extending from the first conduit 110 downstream (relative to the airflow 91 from the compressor section 21 to the turbine section 27) of the first heat exchanger 141. The second conduit 120 extends 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, regulate, or otherwise vary an amount of the airflow 91 from the first conduit 110 through the second conduit 120.

[0078] 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 airflow 91, schematically depicted via arrow 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 airflow 192 to the second location 272 at the turbine section 27. In this manner, it will be appreciated that for the depicted embodiment, the flow control device 130 is positioned between the inlet portion 121 and the outlet portion 122 of the second conduit 120.

[0079] The flow control device 130 can be a valve or any suitable device for regulating, directing, controlling, or otherwise modulating the amount of fluid flow through a passageway or flow path. The flow control device 130 can include an actuated or automatic valve driven by an electrical energy source, a pneumatic energy source (e.g., air, or in particular, at least a portion of the air flow 91), or a fluid source (e.g., liquid and / or gaseous fuel, hydraulic fluid, lubricant, or combinations thereof). The flow control device 130 can include a ball valve, a spool valve, or other suitable type of valve or flow regulating device in accordance with the embodiments depicted and described herein. Thus, the flow control device 130 is configured to modulate the amount of fluid flow through the outlet portion 122 of the second conduit 120, such as schematically depicted via arrows 94.

[0080] In particular embodiments, 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. Thus, the flow control device 130 can be a three-way valve configured to selectively vary the amount of air flow 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. Thus, the flow control device 130 can be configured to modulate the amount of air flow 192 through the outlet portion 122 of the second conduit 120, such as schematically depicted via arrows 194, and further to modulate the egress of at least a portion of the air flow 192 through the third conduit 123, such as schematically depicted via arrows 195. The third conduit 123 can form a bypass passageway to further allow selective regulation, control, or modulation of air flow through the flow control device 130. In particular embodiments, the third conduit 123 allows 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 a third location 273 at the turbine section 27. In certain embodiments, the third location 273 allows bypassing of a gap control system 275 (described below) and allows the air flow 195 to enter the turbine section 27 at the core gas flow path 70 downstream of the gap control system 275, or to mix with the air flow 193 at the turbine frame 308, or to vent to the environment (not depicted).

[0081] Still referring to Figures 4-5 As briefly noted above, the turbine section 27 includes a gap control system 275. An exemplary embodiment of an improved gap control system is depicted in Figures 8-16 , including a housing 300, a manifold assembly 316, and a thermal control ring 314 disposed therein, for example. However, it should be appreciated that Figures 4-5The depicted gap control system 275 can include any suitable structure or component for controlling, regulating, or otherwise modulating the size between the rotor blade tip and the surrounding shroud or wall at the turbine section 27 (also referred to as tip clearance). The gap control system 275 can be an active clearance control (ACC) system configured to dynamically control the tip clearance. In particular, the ACC system can be configured to desirably modulate the tip clearance based on engine operating conditions via the air flow 94 received from the second conduit 120 and provided to the surrounding shroud at the turbine section 27. The volumetric or mass flow rate of the air flow 94 is adjusted or modulated by the flow control device 130. The amount of air flow 94 modulated to the gap control system 275 allows the tip clearance to be desirably adjusted across various engine operating conditions and associated temperature variations at the turbine section 27. As the temperature and rotor speed vary at the turbine section 27 across various engine operating conditions, the flow control device 130 modulates the amount of air flow 94 provided to the gap control system 275 to maintain or provide a desired tip clearance. With respect to a landing-takeoff cycle (LTO) of the engine 10 and aircraft, the engine operating conditions include start-up, idle, takeoff, climb, cruise, approach, or reverse thrust. However, it should be understood that other engine operating conditions and cycles can apply.

[0082] Still referring to Figures 4-5 , the second location 272 at the turbine section 27 is at the gap 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 air flow 94 to the gap control system 275, such as described herein. In particular embodiments, the gap control system 275 is operably coupled to the first turbine assembly or high-speed turbine 28 at the turbine section 27. Accordingly, the engine 10 is configured to receive the air flow 91 from the compressor section 21 and provide a portion of the air flow 94 (from the air flow 91) to the gap control system 275 at the high-speed turbine 28 via the second conduit 120.

[0083] In still particular embodiments, the first conduit 110 is fluidly coupled to a turbine frame 308 positioned between the first turbine assembly or high speed turbine 28 and the second turbine assembly or low speed turbine 30. The turbine frame 308 can include a plurality of vanes 310 in a 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 these embodiments, the first conduit 110 is configured to provide at least a portion of the airflow 91 to the turbine frame 308 at the first location 271. In particular embodiments, the illustrative arrow 193 depicts a portion of the airflow at the first conduit 110 downstream of the junction with the second conduit 120. The airflow 193 is provided to the turbine frame 308 via the first conduit 110. In relation to Figures 8-16 In particular embodiments further depicted and described, the airflow 193 can be provided to the housing 300 and through the plurality of vanes 310 at the turbine frame 308, such as illustratively depicted via arrow 99.

[0084] Referring to Figures 4-5 The turbine frame 308 can include or form one or more channels 309 configured to provide fluid communication of the airflow 193 to the bearing assembly 200. The airflow 193 can provide a cushion fluid for operation of the bearing assembly 200. The cushion fluid can desirably control or dampen vibrations at the bearing assembly 200 or rotor coupled thereto, or allow or generate a desired clearance or vibrational response.

[0085] Now referring specifically to Figure 5 In particular embodiments, the engine 10 includes a second heat exchanger 142 in thermal communication with the airflow at the bypass airflow passage 48. The second heat exchanger 142 can be configured as a surface heat exchanger configured to receive heat or thermal energy from the airflow 194 at the second conduit 120 downstream of the flow control device 130. The heat transfer fluid at the second heat exchanger 142 is the airflow through the bypass airflow passage 48 of the engine 10, such as illustratively depicted via arrow 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 airflow 194 at the second conduit 120 in thermal communication with the bypass airflow 177 at the bypass airflow passage 48. In particular embodiments, 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.

[0086] Referring back generally to Figures 4-5Both, in particular embodiments, the first conduit 110 includes an inlet manifold 111 configured to receive the air stream 91 from the circumferential compressor location 211 at the compressor section 21. It should be appreciated that while the depicted embodiment depicts a single circumferential compressor location 211, the inlet manifold can be configured to receive air streams 91 from multiple circumferential compressor locations 211. Figures 4-5 The depicted embodiment in

[0087] Reference is now made to Figure 6 , providing a perspective view of an embodiment of a portion of an engine 10 according to one or more of Figures 1 to 3 . Figure 6 The embodiments provided in Figures 4-5 may be configured substantially similarly to that described with respect to the embodiments in Figure 6 , the engine 10 can include multiple inlet manifolds 111 evenly spaced or asymmetrically spaced in the circumferential direction C around the compressor section 21. In various embodiments, the multiple inlet manifolds 111 include two (2) or more inlet manifolds. In one embodiment, the multiple inlet manifolds 111 include three (3) inlet manifolds. In another embodiment, the multiple inlet manifolds 111 include four (4) inlet manifolds, and up to 30 inlet manifolds 111.

[0088] In Figure 6 , the first conduit 110 includes a collector 115 configured to receive the air stream 91 from the inlet manifold 111. In particular embodiments, the multiple inlet manifolds 111 are fluidly coupled to a single collector 115 to provide the collected or unified air stream 91 to the first heat exchanger 141. The collector 115 can provide the air stream 91 to the first heat exchanger 141, such as described herein.

[0089] In yet another particular embodiment, the first conduit 110 includes an outlet manifold 112 configured to fluidly communicate the airflow 91 from the first heat exchanger 141 to the turbine section 27 at a first turbine location 271 at the turbine section 27. The engine 10 can include a plurality of outlet manifolds 112 evenly spaced or asymmetrically spaced in 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 extends fluidly from one or more of the plurality of outlet manifolds 112 of the first conduit 110. Accordingly, the plurality of outlet manifolds 112 can extend to a plurality of first turbine locations 271 at different circumferential locations at the turbine section 27.

[0090] It should be appreciated that while the embodiments depicted in FIGS. 1-3 depict a single circumferential first turbine location 271, the first turbine location 271 can include a plurality of circumferential first turbine locations 271. Figures 4-5

[0091] Figures 4-5 The embodiments of the engine 10 provided in FIGS. 1-3 can include the first conduit 110 as a fixed area flow path from the compressor section 21 to the turbine section 27. In other words, the first conduit 110 can include various cross-sectional areas or converging and diverging flow paths. However, the first conduit 110 and the circumferential compressor location 211 can define a fixed or non-linkable flow path area. Such a fixed area flow path allows for a constant volume or mass flow rate of the airflow 91 from the compressor section 21 through the first conduit 110 with respect to a corresponding engine operating condition. In other words, the fixed area flow path allows the first conduit 110 to receive a corresponding flow rate of the airflow 91 with respect to a particular engine operating condition. Accordingly, the embodiments of the engine 10 provided herein allow for a constant airflow 91 to be in thermal communication with the heat transfer fluid flow 221 at the first heat exchanger 141. For example, the flow rate of the heat transfer fluid 221, such as a fuel flow rate or a lubricant flow rate, can be controlled via a schedule, table, graph, or curve that indicates the flow rate versus the engine operating condition. In one embodiment, the airflow 91 at the first conduit 110 can be substantially fixed as a ratio or proportion of the total airflow into the core engine inlet 20 into the compressor section 21. In another embodiment, the airflow 91 at the first conduit 110 can be substantially fixed as a ratio or proportion of the airflow from the low speed compressor 22 into the high speed compressor 24.

[0092] ​Via flow control device 130, engine 10 can specifically include a variable area flow path at second conduit 120. Accordingly, engine 10 can allow for a fixed air flow 193 to turbine frame 308, such as for bearing assembly 200, and a variable air flow 194 to gap control system 275. Flow control device 130 can adjust, link, or otherwise regulate air flow 194 to gap control system 275 as a function of engine operating conditions. Regulation of air flow 194 via flow control device 130 can be a function of inlet air speed (into turbomachine 14 via inlet 20), or inlet air pressure (e.g., corresponding to an altitude of engine 10 during operation or at one or more of the above-described engine operating conditions), or inlet air temperature, or a combination thereof. Additionally or alternatively, regulation of air flow 194 via flow control device 130 can be a function of functional tip clearance at turbine section 27, or a predetermined schedule corresponding to wear or degradation at turbine section 27.

[0093] Certain embodiments of engine 10 include a particular placement of circumferential compressor location 211 at a particular axial stage or other location at compressor section 21 corresponding to a particular pressure range of air flow 91 during operation of engine 10. In various embodiments, circumferential compressor location 211 from which air flow 91 from core gas flow path 70 is received corresponds to a compressor location having air flow therethrough at a pressure between approximately 20 pounds per square inch (psi) and approximately 60 psi, corresponding to between approximately 55% and approximately 75% of an operating envelope, during engine operating conditions. In another embodiment, circumferential compressor location 211 from which air flow 91 from core gas flow path 70 is received can correspond to a compressor location having air flow therethrough at a pressure between approximately 30 pounds per square inch (psi) and approximately 50 psi, during engine operating conditions such as described herein.

[0094] Accordingly, embodiments of the engine 10 provided herein allow the clearance control system 275 and bearing assembly 200 to operate and receive air from the compressor section 21. In certain embodiments, the engine 10 provided herein allows the clearance control system 275 to receive the air flow 91 from the compressor section 21 and not from the bypass airflow passage 48. Additionally or alternatively, the engine 10 provided herein allows the air flow 91 to be received from upstream, forward, or a lower pressure stage of the compressor section as compared to other compressor discharge systems that can receive high energy air from downstream, aft, or a higher pressure stage of the compressor section. Certain of these other compressor discharge systems can 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 the constant air flow 91 through the first conduit 110 to maintain a purge and backflow margin at the turbine frame 308 and bearing assembly 200.

[0095] Referring now to Figures 7A-7B , a flowchart is provided that outlines steps of a method 1000 for operating an engine. The steps of the method 1000 can be stored as instructions and / or executed as operations by embodiments of the engine 10 and computing system 1210 provided herein. Accordingly, the method 1000 can be a computer-implemented method, where one or more steps are stored as instructions at a 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 can cause embodiments of an engine such as described herein with respect to Figures 1-6 to perform operations such as outlined in the flowchart in Figures 7A-7B and further described herein with respect to the method 1000.

[0096] Referring to Figures 7A-7B and to any one or more embodiments depicted in Figures 1-6 , the method 1000 includes, at 1010, initiating rotation of one or both of the high speed spool or the low speed spool, for example to 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 a 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 air flow through the core gas flow path 70 into the combustion section 26 for mixing with liquid and / or gaseous fuel prior to ignition to generate combustion gases.

[0097] The method 1000 further includes, at 1020, compressing the air flow by the compressor section. During operation of the engine 10, the air flow 171 is received at the fan section 16. A portion of the air flow 171 enters the turbine 14 through the core engine inlet 20, such as schematically depicted via arrow 172. The air flow 172 is pressurized by successive rows or stages of compressor blades at the compressor section 21. In particular, the low speed compressor 22 can include a low pressure compressor or supercharger relative to the high speed compressor 24, which includes a high pressure compressor. In certain embodiments, a portion of the air flow 172 compressed by the low speed compressor 22 can be bled or redirected from the core gas flow path 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 air flow 172 and further compresses the air flow, such as schematically depicted via arrow 173 in Figures 1-3

[0098] The method 1000 includes, at 1030, extracting a portion of the compressed air flow from the compressor section, such as described above. The method 1000 can particularly include, at 1030, extracting the portion of the compressed air flow into the first conduit and bypassing the combustion section, such as provided above with respect to the first conduit 110. The method 1000 includes, at 1040, flowing the extracted portion of the compressed air flow through the first conduit (e.g., the first conduit 110) to the turbine section. In particular embodiments, the first conduit bypasses the combustion section when flowing the extracted portion of the compressed air flow to the turbine section. With respect to Figures 1-6 , the portion of the air flow at the compressor section 21 is bled or removed from the core gas flow path 70 and provided to the first conduit 110, such as schematically depicted via arrow 91 in Figures 1-5 The particular embodiments depicted herein can receive the air flow 91 from the compressed air flow 173, 174 from the high speed compressor 24. In still other embodiments, the air flow 91 can be received from the compressed air flow 172 from the low speed compressor 22.

[0099] ​It should be appreciated that the embodiments of engine 10 provided herein advantageously receive a relatively lower pressure and lower temperature flow of air from compressor section 21 and can further avoid structure, complexity, 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 air flow path. Further, it should be appreciated that while specific operating conditions and operating envelopes are provided herein, the engine 10 and / or method 1000 provided herein allow for one or more steps to be performed at any engine operating condition, including up to 100% of total power output. However, specific advantages and benefits are provided herein with respect to engine operation at engine operating conditions that define a majority of the operating envelope. As such, the methods and structures provided herein allow for improved efficiency and reduced fuel consumption.

[0100] In various embodiments, the method 1000 includes, at 1030, extracting a portion of the compressed air flow when the compressed air flow at the compressor section is between approximately 20 pounds per square inch (psi) and approximately 60 psi. In particular embodiments, the method 1000 includes, at 1030, extracting a portion of the compressed air flow when the compressed air flow at the compressor section is between approximately 30 psi and approximately 50 psi. In particular embodiments, the method 1000 includes, at 1035, receiving the portion of the compressed air flow from the compressor section, where the portion of the compressed air flow is between approximately 20 psi and approximately 60 psi, or between approximately 30 psi and approximately 50 psi. In yet further particular embodiments, the method 1000 at 1030 and / or 1035 is performed continuously or constantly with respect to discrete engine operating conditions, such as to allow for a fixed air flow with respect to discrete engine operating conditions.

[0101] In yet another particular embodiment, the method 1000 includes, at 1028, operating the engine at an engine condition between approximately 55% and approximately 75% of the operating envelope or between approximately 60% and approximately 70% of the operating envelope, such as described above. In certain embodiments, one or both of the steps in the method 1000 at 1030 and 1035 occur prior to or concurrently with 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 engine's total power output (e.g., rated thrust), such as described above. In yet another particular embodiment, the method 1000 includes, at 1029, operating the engine between approximately 80% and approximately 88% of the engine's total power output. In certain embodiments, one or more ranges provided herein can define discrete engine operating conditions at which the method 1000 at 1030 and / or 1035 is continuously or constantly performed. In still particular embodiments, the method 1000 includes performing the steps at 1028 and 1029 concurrently.

[0102] The method 1000 can include, at 1050, flowing, via the fluid system, a heat transfer fluid in thermal communication with the extracted portion of the compressed air stream, such as described above. In particular embodiments, Figures 4-5 The fluid system 220 depicted in FIG. 2 is a liquid and / or gaseous fuel system configured to provide a flow of liquid and / or gaseous fuel to the compressed air stream 174 to generate the combustion gas 175. In such embodiments, the fuel is a heat transfer fluid 221 that is in thermal communication with the air stream 91 via the first heat exchanger 141. The fuel stream receives heat or thermal energy from the relatively hot air stream 91, which can advantageously alter certain properties of the fuel, such as viscosity, density, or other properties that can desirably affect combustion, fuel-air mixing, swirl, emissions generation, vibration, or smoke and particulate generation.

[0103] In certain embodiments, the method 1000 can further include flowing, via the fluid system, a plurality of heat transfer fluids in thermal communication with the extracted portion of the compressed air stream. In various embodiments, the method 1000 includes providing one or more of a fuel, a lubricant, a hydraulic fluid, a refrigerant, a supercritical fluid, or another air stream at the heat transfer fluid in thermal communication with the extracted air stream.

[0104] The method 1000 can further include regulating the heat transfer fluid flow to control a temperature of the extracted air stream (e.g., the air stream 91). Regulating the heat transfer fluid flow can include regulating a mass or volumetric flow rate, a pressure, or a temperature of the heat transfer fluid provided in thermal communication with the extracted air stream.

[0105] As provided above, the liquid and / or gaseous fuel streams are mixed with 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, in particular, 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.

[0106] It should be appreciated that the combustion gases 175 release a variable amount of heat at the turbine section 27 based on engine operating conditions. Accordingly, the heat release and turbine rotor speeds can vary the tip gap between the turbine rotor blade tips and surrounding shrouds, such as further described below. It should be appreciated that improved aerodynamics and operating efficiency is generally achieved by minimizing the tip gap. Accordingly, a gap control system is used to regulate the tip gap based on engine operating conditions to improve engine efficiency and performance.

[0107] The method 1000 can further include, at 1060, selectively flowing a portion of the air stream through a second conduit (e.g., the second conduit 120), the second conduit (e.g., the second conduit 120) extending from the first conduit (e.g., the first conduit 110) downstream of the heat exchanger (e.g., the first heat exchanger 141). In particular embodiments, the method 1000 includes, at 1062, varying or regulating, via a flow control device (e.g., the flow control device 130) at the second conduit extending from the first conduit, the portion of the air stream extracted from the first conduit (e.g., the first conduit 110) to the second conduit (e.g., the second conduit 120) downstream of the heat exchanger (e.g., the first heat exchanger 141). In yet other particular embodiments, the method 1000 includes, at 1063, regulating, via the flow control device, a second portion of the air stream extracted from the first conduit to a third conduit extending from the flow control device, such as in the example of FIG. 1. Figures 4-5As depicted by arrow 195. In yet another specific embodiment, method 1000 at 1060 is performed simultaneously with method 1000 at one or more of steps 1028, 1030, or 1035. Thus, method 1000 can allow a continuous, constant, or fixed airflow from the compressor section through the first duct while regulating or altering the airflow through the second duct. In a specific embodiment, method 1000 allows a continuous, constant, or fixed airflow from the compressor section through the first duct and to the turbine section, or in particular, to the bearing assembly, while regulating or altering the airflow through the second duct to the clearance control system. In this way, regulating the airflow through the second duct allows variable airflow to the clearance control system (e.g., clearance control system 275) regardless of whether the engine operating conditions are steady-state (e.g., non-transient or non-changing) or transient (e.g., changing).

[0108] Method 1000 may further include, at 1070, at the clearance control system, selectively changing, altering, or regulating the tip clearance based on the airflow received from the second duct via steps 1060 and / or 1062. It should be understood that the method 1000 provided herein may further provide a method for operating the clearance control system and the bearing assembly. This method may allow for variable flow rate, temperature, pressure, or other physical properties of the airflow through the second duct to the clearance control system, while allowing for a substantially constant or continuous airflow through the first duct relative to engine operating conditions.

[0109] Although Figures 7A-7B Not depicted, but method 1000 may further include generating a bypass airflow through a bypass airflow channel. A portion of the airflow 171 passes through multiple fan blades 40 and around the turbine 14, such as via... Figures 1-3 Arrow 176 in the diagram illustrates this. The airflow 176 entering the bypass airflow passage 48, schematically depicted via arrow 177, is larger in volume or mass and colder than the airflow pressurized by the compressor section 21 within the turbine 14. This can be applied to... Figures 1-3 Any of the embodiments of engine 10 Figure 5 Specifically, a relatively cool bypass airflow 177 thermally connected to airflow 194 via a second heat exchanger 142 is depicted. Thus, method 1000 may further include, at 1064, thermally connecting the bypass airflow to a portion of the airflow extracted into the second duct via the second heat exchanger.

[0110] about Figure 5 The embodiments of engine 10 depicted and described may allow an increase in heat transfer from airflow 194, such as via bypass airflow 177 at bypass airflow passage 48. Furthermore, Figure 5The depicted embodiments, when applied to a turbine engine such as Figure 2 When applied to an open rotor configuration such as described herein, the method 1000 can provide a method for operating an open rotor engine, or in particular, a method for operating a clearance control system of an open rotor engine, or more particularly, a method for operating a clearance control system and bearing assembly of an open rotor engine.

[0111] Reference is now made to Figure 8 , providing an enlarged cross-sectional view of a turbine section portion of a turbomachine 14 in accordance with one or more of Figures 1-3 As shown in Figure 8 , a first turbine assembly is formed by a high speed turbine 28. A first stage 50 of the first turbine assembly includes a plurality of first turbine rotor blades 58 extending within a core gas flow path 70, and further includes an annular array of stator vanes 54 (only one shown) axially spaced apart from an annular array of turbine rotor blades 58 (only one shown) at the high speed turbine 28. In particular embodiments, the high speed turbine 28 further includes a last stage 60 including an annular array of stator vanes 64 (only one shown) axially spaced apart from an annular array of turbine rotor blades 68 (only one shown). The turbine rotor blades 58, 68 extend radially outward from the high speed spool 34 Figure 1 , Figure 2 ) and are coupled to the high speed spool 34 Figure 1 , Figure 2 . The stator vanes 54, 64 and turbine rotor blades 58, 68 at least partially define the core gas flow path 70 for directing combustion gases from the combustion section 26 Figure 1 , Figure 2 through the high speed turbine 28.

[0112] As shown in Figure 8As further shown, the high speed turbine 28 can include one or more shroud assemblies, each forming an annular ring around the annular array of rotor blades. For example, a shroud assembly 72 can 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. Generally, the shroud assembly 72 is radially spaced from the blade tips 76, 78 of each rotor blade 58, 68. A radial or clearance spacing CL is defined between the blade tips 76, 78 and the respective inner surface of the shroud segment 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 a thermal control ring 314 that helps to control thermal growth of the shroud, thereby controlling the radial deflection or clearance spacing CL. A gap control system 275 is used to actively control thermal growth in the shroud assembly. The gap control system 275 is used to minimize the radial blade tip gap CL between the outer blade tip and the shroud, particularly during engine cruise operation, such as described herein.

[0113] 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 can be rotationally separate from the first turbine assembly, such as described with respect to the high speed turbine 28 and low speed turbine 30 described above with reference to FIGS. 1-3. Figure 1

[0114] The casing 300 surrounds the high speed turbine 28. The casing 300 includes a plurality of vanes 310 that extend 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 that extends around the shroud assembly 72 and in a circumferential direction C relative to the centerline axis 12 Figures 1-3 ) of the engine 10. The outer casing wall 312 extends in an axial direction A forward of the rotor blades 58 of the first stage 50 (also referred to as first stage rotor blades 58) of the high speed turbine 28 and aft of the rotor blades 68 of the second or last stage 60 (also referred to as second stage rotor blades 68) of the high speed turbine 28.

[0115] The plurality of vanes 310 extend from the outer casing wall 312. The plurality of vanes 310 extend into the core gas flowpath 70. In certain embodiments described further herein, one or more of the plurality of vanes 310 can be hollow, or include a conduit or passage that allows fluid flow within the vane. The outer casing wall 312 of the casing 300 extends in the axial direction A from a downstream end or trailing edge of the last stage of 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 Figure 8 FIG. 4.​

[0116] It should be understood that conventional turbine housings include separable or engaging flanges, such as bolted flanges or welded flanges, between a high-speed turbine housing and a downstream housing (such as an inter-turbine frame, intermediate turbine frame, medium-speed turbine housing, or low-speed turbine housing). Embodiments of housing 300 provided herein include a single integral structure, such as formed by one or more additive manufacturing processes. Embodiments provided herein further form a single, continuous, and consistent structure, allowing for a single integral extension such as housing 300 provided herein, or further including one or more features integrally formed to housing 300 provided herein.

[0117] The plurality of walls forming the thermal control ring 314 extend in the circumferential direction C and extend outward from the housing wall 312 in the radial direction R. In various embodiments, the thermal control ring 314 includes a front thermal control ring 3141, which is positioned outwardly from the first-stage rotor blade 58 of the high-speed turbine 28, or particularly from the blade tip 76 of the rotor blade 58 of the high-speed turbine 28, in the radial direction R. In some embodiments, such as in Figure 8 As depicted, the front thermal control ring 3141 is positioned along the axial direction A aligned with the first-stage rotor blade 58 (overlapping axial position). In another particular embodiment, the thermal control ring 314 includes a rear thermal control ring 3142, which is positioned radially R outward from the rotor blade 68 of the last stage 60 of the high-speed turbine 28, or particularly from the blade tip 78 of the rotor blade 68 of the high-speed turbine 28. In some embodiments, such as in Figure 8 As depicted, the rear thermal control ring 3142 is positioned along the axial direction A to align with the rotor blades 68 of the last stage 60 of the high-speed turbine 28 (overlapping axial position).

[0118] A front thermal control ring 3141 and a rear thermal control ring 3142 are provided to more effectively control the blade tip clearance CL. Figure 8 As shown in the diagram, it features minimal time lag and thermal control airflow (depending on cooling or heating conditions). The front thermal control ring 3141 and rear thermal control ring 3142 are formed integrally with the housing wall 312 as a single, unified structure of the housing 300. The thermal control ring 314 provides thermal control mass to more effectively adjust the blade tip clearance CL by moving the shroud section 77 radially R. This clearance control can provide a lower operating fuel consumption rate (SFC).

[0119] The integral one-piece construction of the thermal control ring 314 and the outer casing wall 312, particularly where the outer casing wall extends aft of the second or last stage rotor blades 68 of the high speed turbine 28, can allow for improved clearance control, improved thermal control, and improved cooling flow. The structures provided herein allow for the thermal control ring 314 to be positioned radially outward of and axially aligned 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 the exclusion of a flange between the high speed turbine and intermediate turbine frames, intermediate turbine frames between the high speed turbine and a downstream low speed turbine (or such as a mid speed turbine as described herein).

[0120] The embodiments of the integral casing provided herein are generally produced through one or more additive manufacturing processes as described above. While additive manufacturing can generally be applied to form various structures or integrate various components, it should be appreciated that the combination of the integral structures provided herein can overcome issues associated with integral structures while providing unexpected benefits. In one instance, an axially extending casing can generally be susceptible to thermal distortion that can ovalize the core flow passage, which can adversely affect rotor operation as the rotor can rub within a non-concentric flow path. As such, the simple integration of a relatively hot casing around a high speed turbine with a relatively cool casing around an adjacent low speed turbine downstream blade can adversely affect overall engine operation. In another instance, such a large axially extending mass can require additional cooling flow, which can result in increased fuel consumption and overall loss of engine performance.

[0121] The embodiments of the engine provided herein overcome these issues at least in part by positioning a thermal control ring axially aligned with and radially outward of a respective stage of high speed turbine blades. The removal of a flange between a casing around a high speed turbine rotor and a blade casing or frame downstream of the high speed turbine allows for the thermal control ring to be advantageously positioned as disclosed herein.

[0122] Other embodiments of the engine provided herein overcome these issues at least in part through improved cooling flow structures, passages, and conduits. In various embodiments, a manifold assembly 316 surrounds the thermal control ring 314 in the circumferential direction C and the axial direction A. The manifold assembly 316 is configured to provide a fluid flow, such as the air flow 192 from the compressor section 21 (such as described with respect to Figures 4-5 depicted and described) to the thermal control ring 314.

[0123] Still referring to Figure 8 and now also referring to Figures 9-11 and Figure 14 further exemplary embodiments are provided. Figure 8 、 Figure 9 andFigure 14 The embodiments depicted therein can be similarly configured to one another, such as further described below. Figures 9-11 Views of fluid flow and openings at various cross-sections of embodiments of the engine 10 at different circumferential locations of the engine 10 are provided. Each embodiment can be formed via one or more manufacturing methods known in the art. In Figure 14 In particular, the embodiments provided can include a double-walled structure that can be formed via an additive manufacturing process. Various embodiments provided herein can be formed as a unitary single structure, such as via an additive manufacturing process or other suitable manufacturing process.

[0124] Referring to Figures 8-11 and Figure 14 In various embodiments depicted therein, the manifold assembly 316 extends forward and aft of the plurality of axially spaced stages of the plurality of walls forming the thermal control ring 314 in the axial direction A. In particular embodiments, such as in Figure 14 In various embodiments depicted therein, the manifold assembly 316 extends aft in the axial direction A of the plurality of vanes 310. In various embodiments, such as in Figure 8 In exemplary embodiments of the Figure 8 In exemplary embodiments of the

[0125] With particular reference to Figures 9-10 Certain embodiments of the housing 300 include a corrugated feature 399. The corrugated feature 399 includes a shape defining ridges or grooves configured to mitigate the formation of thermal expansion stresses at the housing 300. In certain embodiments, the corrugated feature 399 is formed at the manifold assembly 316. In yet another particular embodiment, the corrugated feature 399 can be formed at the inner manifold 1316 or the outer manifold 2316. The corrugated feature 399 can allow for the unitary single formation of the manifold assembly 316 with the outer housing wall 312, such as described in various embodiments herein.

[0126] Reference is now briefly made to Figure 15The manifold assembly 316 includes a plurality of openings 318 that surround the plurality of walls that form the thermal control ring 314 at the housing 300. The plurality of openings 318 allow fluid flow, schematically depicted via arrow 91, to be in thermal communication with the thermal control ring 314 for the desired heat transfer effect. In various embodiments, the plurality of openings 318 includes an inlet opening 3181 that is configured to allow air flow 91 to enter a first cavity 1321 in thermal communication with the thermal control ring 314, as further described below. The plurality of openings 318 can further include an outlet opening 3182 that is configured to allow at least a portion of the air flow 91 to exit the first cavity 1321 and enter an inner wall conduit 1326, such as further described below, via air flow 92.

[0127] The inlet opening wall 381 extends between an outer portion 346 and an inner portion 347 of the double-walled structure formed by the inner manifold 1316. The inlet opening wall 381 forms an inlet opening flow path 382 that extends through the double-walled structure that is fluidly separate from the inner wall conduit 1326. The inlet opening 3181 and the inlet opening wall 381 allow air flow 91 to pass from the conduit 1324 around the inner manifold 1316 into an air chamber 383 formed between adjacent thermal control rings 314. In particular, the inlet opening wall 381 extends between the outer portion 346 and the inner portion 347 of the inner manifold 1316. The inlet opening flow path 382 formed by the inlet opening wall 381 allows air flow 91 to enter the air chamber 383 while being fluidly isolated from air flow 92 by the inner wall conduit 1326.

[0128] With particular reference back to Figures 9-10 As discussed, the manifold assembly 316 includes an inner manifold 1316 that surrounds the thermal control ring 314 in the circumferential direction C and the axial direction A. The depicted manifold assembly 316 further includes an outer manifold 2316 that surrounds the inner manifold 1316, as discussed above. The channel wall 1318 extends from the inner manifold 1316 to the outer manifold 2316 to form a channel 1320 within the channel wall 1318.

[0129] In certain embodiments, such as Figure 8 As depicted in FIG. 13, the outer manifold 2316 of the manifold assembly 316 extends in the axial direction A at or aft of the plurality of vanes 310. The outer manifold 2316 is further connected to the outer housing wall 312 at or aft of the plurality of vanes 310. In still certain embodiments, such as Figures 9-11As depicted in FIG. 13, the inner manifold 1316 extends to a forward position (terminating forward of the plurality of vanes 310) along the axial direction A of the plurality of vanes 310. The inner manifold 1316 also extends to a rearward position along the axial direction A of the plurality of walls forming the thermal control ring 314. In this manner, the inner manifold 1316 is connected to the outer housing wall 312 forward of the plurality of vanes 310 and rearward of the thermal control ring 314.

[0130] The above discussion with respect to Figure 15 (both in Figures 9-11 FIG. 13) the first cavity 1321 is formed between the inner manifold 1316 and the outer housing wall 312. The thermal control ring 314 is surrounded by the inner manifold 1316 at a location within the first cavity 1321 between the inner manifold 1316 and the outer housing wall 312. The passage 1320 allows fluid communication with the first cavity 1321 between the inner manifold 1316 and the outer housing wall 312. The passage 1320 further allows the air flow 91 to enter into thermal communication with the thermal control ring 314.

[0131] 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 the passage 1320 by the passage wall 1318. In particular embodiments, the passage wall 1318 extends from the outer manifold 2316 to the inner manifold 1316 through the conduit 1324.

[0132] With particular reference to Figures 9-11 , and further with respect to Figure 14 , the conduit 1324 further extends in fluid communication through one or more of the plurality of vanes 310. Figure 10 and Figure 14 The air flow 91 in thermal and fluid communication with the thermal control ring 314 in the first cavity 1321 is particularly depicted. Figure 10 The air flow 91 in thermal and fluid communication with the thermal control ring 314 in the first cavity 1321 is particularly depicted. In various embodiments, the first cavity 1321 is formed to direct the fluid flow directly to the thermal contact portion of the thermal control ring, such as in a perpendicular direction. Figure 11 and Figure 14 The air flow 92 (as the air flow 99 discussed below) is particularly depicted as exiting the first cavity 1321 through the conduit 1324 and subsequently flowing in series through one or more of the plurality of vanes 310. In certain embodiments, the thermal control ring 314 is formed with the outer housing wall 312 to desirably improve gap control. In one embodiment, such as Figure 13B As depicted in FIG. 13, the thermal control ring 314 includes an outer surface extending as a ridge, a groove, or at an acute or zigzag angle (see the more detailed description below).

[0133] Briefly referring to Figure 14 and further depicted in Figure 15 detailed perspective view, in certain embodiments, the inner manifold 1316 is a double-walled structure that forms an inner wall conduit 1326 between the double-walled structure of the inner manifold 1316. The inner wall conduit 1326 can extend in fluid communication to a second cavity 1322 formed between the outer housing wall 312 and the outer wall 170 of the core gas flow path 70. In these embodiments, the integration of the single unitary housing 300, or additionally, the embodiments of the manifold assembly 316, allows for separate flow into the plurality of vanes 310. In particular, the air flow 91 enters the conduit 1324 from the compressor section, such as depicted and described with respect to Figures 1-6 A portion of the air flow 91, depicted via arrow 92, flows into the first cavity 1321 and then into the inner wall conduit 1326 formed at the double-walled structure. The air flow 92 then flows into one or more of the plurality of vanes 310. Additionally, another portion of the air flow 91, depicted via arrow 99, is maintained 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 mixed at the plurality of vanes 310. In other embodiments, the flows 92, 99 are maintained fluidly separated and provided to separate the respective vanes 310, or separate the conduits within each vane 310. Embodiments of the housing 300 and manifold assembly 316, such as providing fluid flow via secondary utilization after being in thermal communication with the thermal control ring 314, rather than outputting the flow to the atmosphere or underhood area of the engine, allow for improved thermal efficiency and improved overall engine efficiency.

[0134] In certain embodiments, the outer wall 170 of the core gas flow path 70 forms an outer shroud segment 77 of the shroud assembly 72. The outer shroud segment 77 is exposed to the core gas flow path 70 and can include a thermal barrier coating or material configured to withstand heat from the combustion gases. The outer shroud segment 77 can be further configured to at least partially rub against one or more stages of vanes at the core gas flow path 70.

[0135] Still referring to Figure 14 and further in Figure 15 , a side view of the housing 300 is provided Figure 15 , and a close-up view of segment A in Figure 16 , and a close-up view of segment A in Figure 16 Figure 17 ​As depicted, the inner manifold 1316 includes a plenum wall 1319 extending from the inner manifold 1316 and surrounding the thermal control ring 314. In certain embodiments, the plenum wall 1319 extends radially inward from the inner portion 347 of the inner manifold 1316. The plenum wall 1319 can be formed as an integral unitary or monolithic structure with the inner manifold 1316 including the outer portion 346 and the inner portion 347. A first cavity 1321 is formed between the outer surface 1325 of the thermal control ring 314 and the plenum wall 1319.

[0136] With particular reference to Figure 16 and 17 , the thermal control ring 314 includes a wall or body 332 extending outward from the outer housing wall 312, such as outward along the radial direction R. In various embodiments, such as described above with respect to the plurality of thermal control rings 314, the body 332 extends substantially annularly along the circumferential direction C Figures 1-3 .

[0137] With more particular reference to Figure 17 , the body 332 forms an internal flow path 330 to allow fluid flow through the thermal control ring 314. Fluid flow through the body 332 allows for variation in temperature or flow rate of the fluid flow through the flow path 330 at the body 332, desirably controlling, varying, or modulating the temperature or thermal gradient at the thermal control ring 314. Further, fluid flow through the body 332 can allow one or more structures attached to or integrally formed to the thermal control ring 314, such as the outer housing wall 312 or the shroud assembly 72, to move based at least in part on the thermal variation provided by the fluid flow, such as to desirably control the clearance spacing CL Figure 8 between the rotor blades 58, 68 and the shroud assembly 72.

[0138] Still with reference to Figure 17 , the depicted exemplary housing 300 further includes a plurality of pins 334 extending from the outer housing wall 312 to the body 332 along the radial direction R of the engine 10 to which the housing 300 is coupled. With brief reference to Figure 18 , a top-down view of the plurality of pins 334 depicts each pin 334. As Figure 17 and 18 illustrate, each pin 334 is spaced apart from one another along the axial direction A of the engine 10 to which the housing 300 is coupled and along the circumferential direction C of the engine 10 to which the housing 300 is coupled Figure 18 . In this manner, adjacent pins 334 define a gap 336 therebetween.

[0139] With particular reference back to Figure 17The flow path 330, which extends radially through the body 332, further extends in fluid communication to a space or gap 336 disposed between the plurality of pins 334. The thermal control ring 314 can form the flow path 330 as a plurality of discrete, circular, or slotted flow paths arranged adjacent in the circumferential direction C. In other embodiments, the thermal control ring 314 forms the flow path 330 as a plurality of arcuate segments extending at least partially in the circumferential direction C. According to any one or more embodiments depicted and described above with respect to Figures 1-15 According to any one or more embodiments depicted and described above with respect to

[0140] During operation, the airflow 91 is received via the gap 336 and through the plurality of pins 334 into the flow path 330 within the body 332. During operation, the airflow 91 proceeds radially through the body 332 and out of the body 332 through an outlet opening 338 at the flow path 330. The outlet opening 338 is formed by the body 332 away from the gap 336 to allow fluid communication from the flow path 330 to the inner wall conduit 1326 formed within the double-walled structure of the inner manifold 1316. According to any one or more embodiments depicted and described above with respect to Figures 1-15 According to any one or more embodiments depicted and described above with respect to

[0141] Still referring to Figure 17 In various embodiments, the 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 can be formed or positioned to contact 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 blocks fluid flow through the first cavity 1321. In particular embodiments, the seal 1323 can form a structural member configured to provide structural support to the inner manifold 1316 and / or the thermal control ring 314. The seal 1323 can further support the body 332 relative to the plurality of pins 334. In certain embodiments, the seal 1323 is a brazing, welding, or other member that attaches 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 can each extend substantially co-directionally with the thermal control ring 314 as a unitary annular component or as a plurality of arcuate segments positioned in a circumferential arrangement.

[0142] In certain embodiments, the outer housing wall 312, the plurality of pins 334, and the body 332 of the thermal control ring 314 are a single unitary structure, such as can be formed by an additive manufacturing process or other suitable manufacturing process. In still certain embodiments, the inner portion 347, the outer portion 346, and the plenum wall 1319 are formed together as a single unitary structure of the inner manifold 1316. In some embodiments, the thermal control ring 314 and the outer housing wall 312 are a single structure separate from the inner manifold 1316. In still some embodiments, the single structure is formed by an additive manufacturing process.

[0143] Reference is now made to Figure 19 , which provides an exemplary embodiment describing operation of the engine 10. Figure 19 The embodiments provided in Figure 16 are configured substantially similar to the embodiments described with respect to Figures 1-6 and Figures 7A-7B . Operation of the system provided herein can be substantially based on that described with respect to the embodiments of the engine 10 described with respect to Figure 19 and Figure 17 . In Figure 17 , the air flow 91 is received at the second location 272, such as through an opening provided by the outer manifold 2316. The air flow 91 is received into the conduit 1324 formed between the outer manifold 2316 and the inner manifold 1316. The air flow 91 is directed into the plenum 383 via the inlet opening 1381 formed by the inner manifold 1316. The air flow 91 is directed through the plurality of pins 334 and through the flow path 330 (see , into the inner wall conduit 1326 (see

[0144] ). Figure 19 In one embodiment, such as depicted in , the air flow 92 can exit from the inner wall conduit 1326 to the outside of the housing 300 or the engine 10 through the opening 1380, such as via the arrow 93. The air flow 93 can cause heat or thermal energy from the thermal control ring 314 to exit to the atmospheric conditions, or to the under-housing or under-shroud area.

[0145] Reference is now made to Figure 20 , which provides a perspective view of a portion of the engine 10. Figure 20 The embodiments provided in Figures 16-19 are configured substantially similar to the embodiments described with respect to Figure 20A plurality of discrete flow paths 330 extending through the thermal control ring 314 in an adjacent circumferential arrangement are depicted. A plurality of outlet openings 3182 are formed through the inner portion 347 of the inner manifold 1316 corresponding to the plurality of flow paths 330 and outlet openings 338 at the thermal control ring 314. Accordingly, the engine 10 can form a plurality of flow paths 330 and outlet openings 338 at the thermal control ring 314 in an 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 an arrangement can allow the air flow 92 to exit from within the thermal control ring 314 into the inner wall conduit 1326.

[0146] Reference is now made briefly back to Figure 21 , embodiments are provided Figure 20 depicting a side cross-sectional view of the embodiments provided in Figure 21 The embodiments in

[0147] Reference is now made briefly back to Figure 12 and Figures 13A-13D additional aspects of the present disclosure are described. Figure 12 A partial circumferential view of embodiments of the manifold assembly 316 is provided. Further, Figures 13A-13D a cross-sectional view of the embodiments depicted in Figure 12 is provided (for Figures 13A-13D The reference numerals in Figure 12 are indicated for each of Figure 12 and Figure 13C As previously described, various embodiments of the manifold assembly 316 are formed via one or more additive manufacturing processes. With particular reference to the close-up views of Figure 12 and Figure 13C In various embodiments, the member 3316 extends to the inner manifold 1316 and the outer manifold 2316. The member 3316 extends from the inner manifold 1316 to the outer manifold 2316 at an acute angle (e.g., a V-, Z-, or other angled cross-section). In various embodiments, the member 3316 extends along a first direction, depicted schematically via arrow 95, and along a second direction opposite the first direction, depicted schematically via arrow 96.

[0148] 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, casing 300, and manifold assembly 316 provided herein include an integrated unitary structure, such as a casing extending over multiple stages of a high speed turbine, or further including a turbine inter- frame, or further including all or portions of a manifold, such as can be formed by additive manufacturing processes heretofore not possible or feasible. Embodiments depicted and described herein allow for improved and advantageous positioning of the thermal control ring 314, flow paths 330 therethrough, and 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 exclusion of flanges and subassemblies to be integrated components. The particular combination of these features allows for improved heat transfer properties and reduced thermal gradients. The improved heat transfer properties include, among other things, reduced heat transfer coefficients at certain features, such as the plurality of walls, bodies, pins, and / or flow paths forming the thermal control ring 314, as compared to known clearance control systems. Such improvements can mitigate or eliminate undesirable or excessive distortion, ovalization, bowing, or other changes in geometry of the casing 300 that can adversely affect deflection or result in undesirable contact with turbine rotor blades 58 at the high speed turbine 28.

[0149] Embodiments of the engine 10 and casing 300 provided herein include an integrated unitary casing for a high speed turbine 28, along with a turbine center frame or intermediate turbine frame 308 formed from outer casing walls 312 and a plurality of vanes 310, and positioned downstream along a core gas flow path 70 of the high speed turbine 28 and upstream along a core gas flow path 70 of a low or intermediate pressure turbine, such as depicted at turbine 30. Embodiments provided herein further include an integrated unitary clearance control manifold configured to provide heat transfer fluid to the thermal control ring, for example. The integrated unitary structure can further allow for improved positioning of the thermal control ring relative to the turbine rotor, such as to provide improved clearance control across the turbine rotor assembly.

[0150] It should be understood that the description of the Figures 1-6 The depicted and described conduits 110, 120, 123, flow control devices 130, or heat exchangers 141, 142 can be provided to the casing 300, manifold assembly 316, and regarding Figures 8-21other structures described herein. However, various embodiments of the engine 10 provided herein can include one or more of the conduits 110, 120, 123, flow control devices 130, or heat exchangers 141, 142 that provide an airflow to any suitable gap control system, turbine section, or bearing assembly. Such structures, when combined with any suitable gap control system, turbine section, or bearing assembly, can provide one or more of the advantages and benefits described herein. Alternatively, various embodiments of the engine 10 provided herein can include one or more of the housings 300 or manifold assemblies 316 that receive an airflow from any suitable conduit, passage, flow path, tube, or other structure. Such structures, when combined with any suitable conduit or heat exchanger, can provide one or more of the advantages and benefits described herein. The benefits and advantages described with respect to the conduits, flow control devices, heat exchangers, housings, or manifolds, when combined together, can be additive of the benefits and advantages described herein.

[0151] Embodiments of the conduits 110, 120, 123 and heat exchangers 141, 142 provided herein can be formed at least partially by one or more additive manufacturing processes, such as described herein. For example, the first heat exchanger 141 can be integrally formed with the first conduit 110, or the second heat exchanger 142 can be integrally formed with the second conduit 120 or portions thereof. In another example, all or portions of the first conduit 110 including one or more of the inlet manifolds 111, outlet manifolds 112, or collector 115 can be integrally formed as a single unitary component. In yet another example, all or portions of the second conduit 120 including one or more of the inlet portions 121 or outlet portions 122 can be formed as a single unitary component. Still further, certain combinations of portions of the first conduit 110, the second conduit 120, and the third conduit 123 can be integrally formed with one another. For example, the outlet manifolds 112 can be formed as a single unitary component with the inlet portions 121. In another example, a housing surrounding the compressor section 21 can be integrally formed with the inlet manifolds 111. The collector 115 can be integrally formed with the first heat exchanger 141. The second heat exchanger 142 can be integrally formed with the outlet portions 122.

[0152] 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 present disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent

[0153] Further aspects of the present disclosure are provided by the subject matter of the following clauses:

[0154] A gas turbine engine, wherein the gas turbine engine defines an axial direction, a centerline axis parallel to the axial direction, a radial direction extending 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 extending within a gas flowpath; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises an outer casing wall extending around the first turbine rotor assembly; a plurality of vanes extending from the outer casing wall and extending within the gas flowpath at a location aft of the first turbine rotor assembly; and a thermal control ring positioned outside the outer casing wall along the radial direction, 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.

[0155] The engine of one or more of these clauses, wherein the plurality of pins define a void between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid flow through the void.

[0156] The engine of one or more of these clauses, wherein a flowpath is formed through the body of the thermal control ring.

[0157] The engine of one or more of these clauses, wherein a flowpath extends through the body along the radial direction, and wherein the flowpath provides fluid communication between the void formed by the plurality of pins and an outlet opening positioned opposite the body from the void.

[0158] 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.

[0159] The engine of one or more of these clauses, comprising: an inner manifold forming a double-walled structure comprising an outer portion separate 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 air flow from the flowpath at the thermal control ring into the inner wall conduit.

[0160] 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.

[0161] The engine of one or more of these clauses, wherein the inner manifold is connected to the outer housing wall forward of the plurality of vanes.

[0162] The engine of one or more of these clauses, wherein the inner manifold forms an inlet opening through the double-walled structure.

[0163] The engine of one or more of these clauses, wherein the inner manifold includes an inlet opening wall forming an inlet opening flow path that is fluidly separate from the inner wall conduit.

[0164] 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.

[0165] The engine of one or more of these clauses, wherein a plenum wall extends from the inner manifold and surrounds the thermal control ring.

[0166] 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.

[0167] 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.

[0168] The engine of one or more of these clauses, wherein the inner portion of the inner manifold forms the outlet openings in an adjacent circumferential arrangement radially outward of corresponding outlet openings of the thermal control ring.

[0169] The engine of one or more of these clauses, wherein the inner manifold extends forward along the axial direction of the plurality of vanes, and wherein the inner manifold is connected to the outer housing wall forward of the plurality of vanes.

[0170] The engine of one or more of these clauses, wherein the outer housing wall, the plurality of pins, and the body of the thermal control ring are a single unitary structure.

[0171] A casing for a gas turbine engine defining an axial direction, a radial direction, a circumferential direction, and a gas flow path, the gas turbine engine including a first turbine rotor assembly including a plurality of first turbine rotor blades extending within the gas flow path, the casing including an outer casing wall configured to extend around the first turbine rotor assembly when the casing is installed in the gas turbine engine, a plurality of vanes extending from the outer casing wall when the casing is installed in the gas turbine engine and configured to extend into the gas flow path at a location aft of the first turbine rotor assembly, and a thermal control ring positioned outward of the outer casing wall along the radial direction, and wherein the thermal control ring includes a body and a plurality of pins, and wherein the plurality of pins extend between the outer casing wall and the body.

[0172] The casing according to one or more of the clauses, wherein the plurality of pins define a void between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid flow through the void.

[0173] The casing according to one or more of the clauses, wherein a flow path is formed through the body of the thermal control ring, and wherein the flow path provides fluid communication between the void formed by the plurality of pins and an outlet opening positioned opposite the body from the void.

Claims

1. A gas turbine engine, wherein the gas turbine engine defines an axial direction, a centerline axis parallel to the axial direction, a radial direction extending from the centerline axis, and a circumferential direction relative to the centerline axis, characterized by, The gas turbine engine comprises: a first turbine rotor assembly comprising a plurality of first turbine rotor blades extending within a gas flow path; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises: an outer casing wall extending around the first turbine rotor assembly; a plurality of vanes extending from the outer casing wall and extending within the gas flow path at a location rearward of the first turbine rotor assembly; and a thermal control ring positioned radially outward of 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; wherein the plurality of pins define a gap between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid flow through the gap, and a flow path is formed through an interior of the body of the thermal control ring.

2. The engine of claim 1, wherein wherein the flow path extends through the body in the radial direction, and wherein the flow path provides fluid communication between the gap formed by the plurality of pins and an outlet opening positioned opposite the body from the gap.

3. The engine of claim 2, wherein wherein the flow path is a plurality of discrete flow paths in adjacent arrangement along the circumferential direction.

4. The engine of claim 2, wherein The engine comprises: an inner manifold forming a double walled 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 air flow from the flow path at the thermal control ring into the inner wall conduit.

5. The engine of claim 4, wherein wherein the inner manifold surrounds the thermal control ring in the circumferential direction and the axial direction.

6. The engine of claim 5, wherein wherein the inner manifold is connected to the outer casing wall forward of the plurality of vanes.

7. The engine of claim 4, wherein wherein the inner manifold forms an inlet opening through the double walled structure.

8. The engine of claim 7, wherein wherein the inner manifold comprises an inlet opening wall forming an inlet opening flow path fluidly separate from the inner wall conduit.

9. The engine of claim 4, wherein, wherein a seal is positioned in contact with the inner portion of the inner manifold and the outer surface of the thermal control ring.

10. The engine of claim 4, wherein, wherein a plenum wall extends from the inner manifold and surrounds the thermal control ring.

11. The engine of claim 10, wherein, wherein a cavity is formed between an outer surface of the thermal control ring and the plenum wall.

12. The engine of claim 11, wherein, wherein a seal is positioned in contact with the outer surface of the thermal control ring and the plenum wall.

13. The engine of claim 4, wherein wherein the inner portion of the inner manifold forms the outlet opening in an adjacent circumferential arrangement radially outward of a corresponding outlet opening of the thermal control ring.

14. The engine of claim 4, wherein, wherein the inner manifold extends forward in 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.

15. The engine of claim 1, wherein, wherein the outer casing wall, the plurality of pins, and the body of the thermal control ring are a single unitary structure.

16. A casing for a gas turbine engine defining an axial direction, a radial direction, a circumferential direction, and a gas flow path, the gas turbine engine including a first turbine rotor assembly including a plurality of first turbine rotor blades extending within the gas flow path, characterized by, The casing comprises: an outer casing wall configured to extend around the first turbine rotor assembly when the casing is installed in the gas turbine engine; a plurality of vanes extending from the outer casing wall and configured to extend into the gas flow path 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 outside the outer casing wall along the radial direction, and wherein the thermal control ring includes a body and a plurality of pins, and wherein the plurality of pins extend between the outer casing wall and the body; wherein the plurality of pins define a void between the outer casing wall and the body of the thermal control ring, wherein the plurality of pins allow fluid to flow through the void, and a flow path is formed through an interior of the body of the thermal control ring.

17. The case of claim 16, wherein, wherein the flow path provides fluid communication between the void formed by the plurality of pins and an outlet opening positioned opposite the body from the void.

Images