Gas turbine engine having a fluid conduit system and method of operating the same

Abstract

A method of operating a gas turbine engine includes extracting an air flow from a compressor section of the gas turbine engine into a first conduit; flowing the extracted air flow through the first conduit to a first location at the turbine section of the turbine section, wherein a second conduit is in fluid communication with the turbine section at a second location; flowing a heat transfer fluid to a first heat exchanger, the first heat exchanger positioned in thermal communication with the air flow through the first conduit, the heat transfer fluid in thermal communication with the extracted air flow through the first conduit via the first heat exchanger; and regulating, via a flow control device, a portion of the air flow extracted from the first conduit to the second conduit downstream of the first heat exchanger.

Description

[0001] Cross-references to related applications

[0002] This application claims priority to Polish patent application number P.439448, 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 fluid conduits for gas turbine engines, such as fluid conduits for clearance control structures in the turbine section of a gas turbine engine. 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. Attached Figure Description

[0008] In the specification with reference to the accompanying drawings, a complete and implementable disclosure, including its best mode, is set forth for those skilled in the art, wherein:

[0009] Figure 1 This is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of this disclosure;

[0010] Figure 2 This is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of this disclosure;

[0011] Figure 3 This is an exemplary schematic cross-sectional view of an embodiment of a gas turbine engine according to aspects of this disclosure;

[0012] Figure 4 This is a schematic cross-sectional view as part of an embodiment of a gas turbine engine according to aspects of this disclosure;

[0013] Figure 5 This is a schematic cross-sectional view as part of an embodiment of a gas turbine engine according to an additional aspect of this disclosure;

[0014] Figure 6 This is a perspective view as part of an embodiment of a gas turbine engine according to aspects of this disclosure;

[0015] Figures 7A-7B A flowchart depicts the steps of a method for operating an engine, outlining aspects of this disclosure;

[0016] Figure 8-11 This is an exemplary schematic cross-sectional view of an embodiment of a turbine section and housing portion according to aspects of this disclosure;

[0017] Figure 12 This is an exemplary perspective view of an embodiment of a portion of a turbine section manifold according to aspects of this disclosure;

[0018] Figures 13A-13D yes Figure 12 An exemplary cross-sectional view of an embodiment of the manifold provided in the document;

[0019] Figure 14 This is an exemplary schematic cross-sectional view of an embodiment of a turbine section and housing portion according to aspects of this disclosure;

[0020] Figure 15 This is an exemplary perspective view of an embodiment of a portion of a turbine section manifold according to aspects of this disclosure;

[0021] Figure 16 This is an exemplary schematic cross-sectional view of an embodiment of a turbine section and housing portion according to aspects of this disclosure;

[0022] Figure 17 Based on the aspects of this disclosure Figure 16 Detailed view of the exemplary schematic cross-sectional view of the embodiment;

[0023] Figure 18 This is a top-down view of an exemplary embodiment of a plurality of pins of a thermal control ring according to aspects of this disclosure;

[0024] Figure 19It is adopted in accordance with the aspects of this disclosure. Figure 16 An exemplary schematic diagram of the airflow in the turbine section and casing;

[0025] Figure 20 It is a perspective view of a portion of the engine according to an aspect of this disclosure; and

[0026] Figure 21 Based on the aspects of this disclosure Figure 20 A cross-sectional view of an embodiment of the engine provided in the document.

[0027] Reference characters used repeatedly in this specification and drawings are intended to indicate the same or similar features or elements of this disclosure. Detailed Implementation

[0028] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to designate features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to designate similar or analogous components of this disclosure.

[0029] The term "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 superior to or advantageous to other implementations. Furthermore, unless explicitly indicated otherwise, all embodiments described herein should be considered exemplary.

[0030] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of an individual component.

[0031] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction in which fluid flows out of the path, and "downstream" refers to the direction in which fluid flows towards the path.

[0032] The singular forms “a,” “a,” and “the” include plural references unless the context clearly indicates otherwise.

[0033] The terms “connection,” “fixation,” “attachment,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0034] The approximate language used throughout this specification and claims is applied to modify any permissible variations that will not result in a change to the essential function associated therewith. Therefore, values ​​modified by one or more terms such as “approximately,” “about,” and “substantially” are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the part and / or system. For example, approximate language may refer to a margin of 1%, 2%, 4%, 10%, 15%, or 20%.

[0035] Throughout this specification and claims, scope limitations are combined and interchanged, and unless the context or language otherwise indicates otherwise, these scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints can be combined independently of each other.

[0036] Pressure values ​​and their ranges are expressed as absolute pressure measurements (psia) or equivalent values. The pressure values ​​and ranges provided herein can be converted to ranges of gauge pressure, or other pressure units, or other units, measurements, or combinations thereof corresponding to the values ​​and / or ranges disclosed herein.

[0037] The term "total power output" refers to the engine's maximum rated power output.

[0038] The term "operational envelope" refers to a cycle, mission, or set of maneuvers in which the engine can operate normally. In one embodiment, a landing-takeoff (LTO) cycle may define the operational envelope. An LTO cycle comprising one or more combinations of engine operating conditions, including start-up, idling, takeoff, cruise, and approach, may collectively define the operational envelope. In various embodiments, the cruise condition defines the majority of the operational envelope, such as defining the majority of the engine operation time or duration. In some embodiments, the cruise condition is between approximately 55% and 75% of the operational envelope. In other words, the cruise condition may define approximately 55% to approximately 75% of the duration of engine operation from start-up to shutdown after the approach operational condition. In another embodiment, the cruise condition may define approximately 60% to approximately 70% of the duration of engine operation.

[0039] 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 the power or thrust condition in between. In some embodiments, the cruise condition corresponds to approximately 75% to approximately 90% of the engine's total power output. In still some embodiments, the cruise condition corresponds to approximately 80% to 88% of the engine's total power output.

[0040] As used herein, "third flow" refers to a non-primary airflow that can increase fluid energy to generate a small amount of total propulsion system thrust. The pressure ratio of the third flow can be higher than that of the primary propulsion flow (e.g., bypass or propeller-driven propulsion flow). Thrust can be generated through dedicated nozzles or by mixing the airflow through the third flow with, for example, the primary propulsion flow or core airflow entering a common nozzle.

[0041] In some exemplary embodiments, the operating temperature of the airflow through the third flow can be less than the engine's maximum compressor discharge temperature, and more specifically, less than 350 degrees Fahrenheit (e.g., less than 300 degrees Fahrenheit, less than 250 degrees Fahrenheit, less than 200 degrees Fahrenheit, and at least as high as the ambient temperature). In some exemplary embodiments, these operating temperatures can facilitate heat transfer to or out of the airflow through the third flow and the separate fluid flow. Furthermore, in some exemplary embodiments, during takeoff, or more specifically, when operating at sea-level rated takeoff power, static flight speed, and an ambient temperature of 86 degrees Fahrenheit, the airflow through the third flow can contribute less than 50% (and at least, for example, 2%) of the total engine thrust.

[0042] Furthermore, in some exemplary embodiments, aspects of the third flow (e.g., airflow, mixing, or exhaust properties) and the aforementioned exemplary percentage contribution to total thrust can be passively adjusted during engine operation or purposefully modified by using engine control features (such as fuel flow, motor power, variable stator, variable inlet guide vanes, valves, variable exhaust port geometry, or fluidity characteristics) to adjust or optimize overall system performance under a wide range of potential operating conditions.

[0043] The term “turbine” or “turbomachinery” refers to a machine that includes one or more compressors, a heating section (e.g., a combustion section), and one or more turbines that together generate torque output.

[0044] The term "gas turbine engine" refers to an engine that has a turbine as its power source, in whole or in part. Examples of gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc.

[0045] The term "combustion section" refers to any heat addition system used in a turbine. For example, the term combustion section can refer to a section that includes one or more of a knock combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In some example embodiments, the combustion section may include an annular burner, a canister burner, an annular tube burner, a vortex burner (TVC), or other suitable combustion systems, or combinations thereof.

[0046] The terms “low” and “high,” or their corresponding comparatives (e.g., lower, higher, where applicable), when used with compressors, turbines, shafts, or spool components, refer respectively to relative speeds within the engine, unless otherwise specified. For example, “low turbine” or “low-speed turbine” defines a component configured to operate at a lower speed (such as the maximum permissible speed) than a “high turbine” or “high-speed turbine” at the engine.

[0047] As used herein, the term "at" refers to the position of a first object relative to a second object (e.g., the first object being located at or positioned at the second object), meaning that the first object is wholly or partially located within the second object, the first object is in contact with the second object, or the first object is positioned as closest to the second object (relative to any other surrounding related component).

[0048] One or more components of the turbine engine described below can be manufactured or formed using any suitable process, such as additive manufacturing or 3D printing. The use of such a process can allow such components to be formed integrally, as a single monolithic part, or as any suitable number of sub-parts. In particular, additive manufacturing processes can allow such components to be formed integrally and include various features that would be impossible to achieve using prior manufacturing methods. For example, the additive manufacturing methods described herein can allow the manufacture of channels, conduits, cavities, openings, housings, manifolds, double walls, heat exchangers, or other components, or the specific positioning and integration of these components with unique features, constructions, thicknesses, materials, densities, fluid passages, manifolds, and mounting structures that might be impossible or impractical using prior manufacturing methods. Some of these features are described herein.

[0049] For example, suitable additive manufacturing techniques according to this disclosure include fused deposition modeling (FDM), selective laser sintering (SLS), 3D printing such as by inkjet, laser jetting and binder jetting, stereolithography (SLA), direct selective laser sintering (DSLS), electron beam sintering (EBS), electron beam melting (EBM), laser engineered net forming (LENS), laser net forming 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.

[0050] Suitable powder materials for manufacturing the structures provided herein as a single, integral structure include metal alloys, polymers, 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 may include those designed to have good oxidation resistance, referred to as “superalloys,” which possess acceptable strength at elevated operating temperatures in gas turbine engines, such as Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-850, ECY 768, 282, X45, PWA 1483, and CMSX (e.g., CMSX-4) single-crystal alloys. The manufactured objects disclosed herein can be formed having one or more selected crystalline microstructures, such as directional solidification (“DS”) or single crystals (“SX”).

[0051] Embodiments of gas turbine engines including improved clearance control systems are provided. The engines reduce weight and the number of pipes, manifolds, or ducts outside the outer core casing or fan casing by reducing or eliminating air drawn from the fan bypass passage for cooling at the turbine section. The embodiments provided herein allow engines without a fan casing, such as open rotor engines or propeller fan engines, to have and operate improved clearance control, cooling systems, or air systems for the turbine section and / or bearing assemblies. It should be understood that while these embodiments can be applied to turbofan engines including nacelles and fan casings, the embodiments provided herein allow engines without a nacelle, fan casing, or other structures surrounding the fan section to receive air for turbine section cooling, clearance control, or bearing assemblies.

[0052] The improved gas turbine engine described herein may additionally or alternatively allow the removal of low-pressure and / or cryogenic air from the compressor section for cooling or clearance control in the turbine section and bearing assemblies. Some clearance control systems can generally utilize high-energy air (i.e., high-pressure and / or high-temperature air) such as from the downstream stage of a high-pressure compressor, and mix it with one or more other air sources such as from other compressor stages or from fan airflow. This high-energy air reduces engine efficiency, such as by removing energy from thermodynamic and combustion processes, or by reducing the required heat load before the air becomes suitable for cooling or clearance control in the turbine section. Furthermore, some clearance control systems may not be suitable for additionally supplying air to the bearing assemblies for cooling, buffering, or other purposes in the bearing assemblies.

[0053] Another aspect of this disclosure relates to an improved turbine housing that allows for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. Embodiments of the engine, housing, and manifold provided herein include a single, integral structure, which can be formed through additive manufacturing processes that have not been possible or feasible until now. Embodiments depicted and described herein allow for improved and advantageous positioning of the thermal control ring for improved clearance control response; improved formation and positioning of openings, channels, and conduits to allow for more efficient utilization and movement of heat transfer fluids; and weight reduction, such as by eliminating flanges and sub-assemblies into a single component. Specific combinations of these features allow for improved heat transfer properties and reduced thermal gradients. Improved heat transfer properties specifically include lower heat transfer coefficients at certain features, such as at the multiple walls forming the thermal control ring as provided herein. This improvement can mitigate or eliminate undesirable or excessive deformation, ellipticing, bending, or other changes in housing geometry that may adversely affect deflection or lead to undesirable contact with the turbine rotor.

[0054] The embodiments provided herein include, for example, a single, integrated high-speed turbine housing, and a turbine center frame or intermediate turbine frame positioned downstream of a high-speed turbine and upstream of a low- or medium-pressure turbine. The embodiments provided herein further include, for example, a single, integrated clearance control manifold configured to provide heat transfer fluid to a thermal control ring. The single, integrated structure can further allow for improved positioning of the thermal control ring relative to the turbine rotor, such as providing improved clearance control across the turbine rotor assembly.

[0055] As used herein, the term "monolithic" to describe a structure refers to a structure formed monolithically from a continuous material or group of materials, without seams, joints, etc. The monolithic structures described herein can be formed to have the structure by additive manufacturing, or alternatively, by casting or the like.

[0056] Now refer to the attached diagram, Figure 1This is a schematic cross-sectional view of an exemplary gas turbine engine 10, referred to herein as "engine 10," which can be incorporated into various embodiments of this disclosure. Specific embodiments of engine 10 may be configured as a turbofan, turboprop, turbine shaft, or propeller-fan gas turbine engine, or as one or more gas turbine engines configured as a hybrid electric gas turbine engine, or other gas turbine engine configurations.

[0057] like Figure 1 As shown, for reference purposes, engine 10 has a longitudinal or axial centerline axis 12 extending therethrough. Generally, engine 10 may include a turbine 14 disposed downstream of fan section 16.

[0058] Engine 10 includes compressor section 21 arranged in series with turbine section 27. Turbine 14 generally includes a substantially tubular outer casing 18 defining an annular inlet 20. Casing 18 may be formed of multiple casings. Casing 18 surrounds compressor section 21, combustion section 26, and turbine section 27 in series flow arrangement. In a particular embodiment, compressor section 21 includes a turbocharger or low-speed compressor 22 and a high-speed compressor 24. In yet another particular embodiment, 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 impeller blades 116 and rotor blades 118). Injector exhaust nozzle section 32 is positioned downstream of turbine section 27. High-speed shaft or spool 34 drivesly connects high-speed turbine 28 to high-speed compressor 24. Low-speed shaft or spool 36 drivesly connects low-speed turbine 30 to low-speed compressor 22. The low-speed spool 36 can also be connected to the fan shaft or spool 38 of the fan section 16. In a particular embodiment, the low-speed spool 36 can be directly connected to the fan spool 38, such as in a direct drive configuration. In alternative configurations, such as... Figure 1 As depicted by the dashed lines, the low-speed spool 36 can be connected to the fan spool 38 via a gear assembly 37, such as to configure the engine 10 in an indirect drive or gear-driven configuration that allows the fan spool 38 to operate at higher or lower speeds than the low-speed spool 36. Such a gear assembly can be included between any suitable shafts / spools within the engine 10 as desired or required.

[0059] 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.

[0060] 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.

[0061] 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.

[0062] 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.

[0063] 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 depicted herein, the bypass airflow passage 48 can generally be included downstream of one or more stages of the plurality of fan blades 40 or the low-speed compressor 22, and bypasses or surrounds at least a portion of the high-speed compressor 24, and has any flow path through which the bypass airflow 177 providing thrust is provided. Therefore, certain embodiments of the engine 10 provided herein can be configured as a third-flow or adaptive cycle engine having a plurality of bypass airflow passages 48 downstream of one or more stages of the plurality of fan blades 40 and / or the low-speed compressor 22 and upstream of at least a portion of the high-speed compressor 24, one or more of the plurality of bypass airflow passages 48 being configured as a “third flow”.

[0064] Engine 10 includes a computing system 1210 configured to operate. The computing system 1210 is communicatively coupled to turbine 14 and / or starter motor (not depicted) to regulate, modulate, maintain, alter, or connect any one or more control surfaces to generate one or more embodiments of airflow, heat transfer fluid flow, and / or liquid and / or gaseous fuel according to aspects of this disclosure provided herein. The computing system 1210 may generally correspond to any suitable processor-based device, including one or more computing devices. Some embodiments of the computing system 1210 include a full-authority digital engine controller (FADEC), a digital engine controller (DEC), or other suitable computing devices configured to operate engine 10.

[0065] The computing system 1210 may include one or more processors 1212 and one or more associated memory devices 1214 configured to perform various computer-implemented functions, such as the steps of the methods described herein. As used herein, the term "processor" refers not only to integrated circuits included in computers in the art, but also to controllers, microcontrollers, microcomputers, programmable logic controllers (PLCs), application-specific integrated circuits (ASICs), field-programmable gate arrays (FPGAs), and other programmable circuits. Additionally, the memory 1214 may generally include memory elements, including but not limited to computer-readable media (e.g., random access memory (RAM)), computer-readable non-volatile media (e.g., flash memory), compact disc-read-only memory (CD-ROM), magneto-optical disk (MOD), digital versatile optical disk (DVD), non-transitory computer-readable media, and / or other suitable memory elements, or combinations thereof.

[0066] The computing system 1210 may include control logic 1216 stored in memory 1214. Control logic 1216 may include computer-readable instructions that, when executed by one or more processors 1212, cause one or more processors 1212 to perform operations such as those outlined in one or more steps of method 1000, which is further provided below. In various other embodiments, memory 1214 may store graphs, tables, functions, lookups, schedules, etc., corresponding to flow rates or speeds, pressures, or temperatures associated with the airflow, heat transfer fluid flow, or fuel flow provided herein. Instructions may be software written in any suitable programming language or may be implemented in hardware. Additionally and / or alternatively, instructions may be executed in logically and / or virtually separate threads on the processor.

[0067] The computing system 1210 may also include a communication interface module 1230. In various embodiments, the communication interface module 1230 may include associated electronic circuitry for transmitting and receiving data. Thus, the communication interface module 1230 of the computing system 1210 may be used to receive data from one or more control surfaces, sensors, measuring devices, or instruments, or calculated or measured values ​​corresponding to one or more parts of the engine 10 provided herein, and may perform one or more steps of the method 1000 provided herein. The computing system 1210 may also include a network interface for, for example, communicating with other components of the engine 10. The network interface may include any suitable components for communicating with one or more network interfaces, including, for example, a transmitter, receiver, port, controller, antenna, and / or other suitable components.

[0068] It should be understood that the communication interface module 1230 can be any combination of suitable wired and / or wireless communication interfaces, and thus can be communicatively coupled to one or more components of the device via wired and / or wireless connections. In this way, 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 may include a SATCOM network, ACARS network, ARINC network, SITA network, AVICOM network, VHF network, HF network, Wi-Fi network, WiMAX network, Gatelink network, etc.

[0069] Now for reference Figure 2 It provides information about Figure 1 An exemplary embodiment of the open rotor configuration of the engine 10 is depicted and described. Figure 2 The embodiment of the engine 10 provided in the middle is configured to be with Figure 1 The basic structure is similar to that provided in [the text]. However, in [the text]... Figure 2In the open rotor configuration of engine 10, there is no fan casing or nacelle 42 surrounding the multiple fan blades 40. Figure 1 (As depicted in the image). A bypass airflow passage 48 is formed downstream of the plurality of fan blades 40, or in particular downstream of the outlet guide vane 44, and radially outward of the outer portion of the turbine 14.

[0070] Now for reference Figure 3 Provided according to Figure 2 An exemplary embodiment of the open rotor structure. Figure 3 The embodiments provided further include a plurality of bypass airflow passages 48 formed downstream of the plurality of fan blades 40, as described above. In a particular embodiment depicted, the engine 10 includes a first bypass airflow passage 48A and a second bypass airflow passage 48B. The second bypass airflow passage 48B extends from a location between the low-speed compressor 22 and the high-speed compressor 24 to an exhaust port leading to the atmosphere (although in other embodiments, the second bypass airflow passage 48B may extend to the first bypass airflow passage 48A). A connecting impeller or door structure 43 may be positioned at the second bypass airflow passage 48B. The door structure 43 may include any suitable type of actuable wall, impeller, door, or other structure configured to desirably alter the airflow 172 received from the core gas flow path 70 and allowed to pass through the second bypass airflow passage 48B (such as schematically depicted via arrow 177). The second bypass airflow passage 48B may be referred to as a third flow.

[0071] Although Figure 3 A three-flow or adaptive cycle engine 10 with an open rotor configuration is depicted; however, it should be understood that the adaptive cycle configuration may also include a nacelle surrounding the fan section, such as regarding... Figure 1 Depicted and described. In this way, it should be further understood that while some of the advantages and benefits provided herein may benefit turbofan engines with nacelles, the embodiments and arrangements of the components provided herein can overcome the problems or challenges peculiar to open rotor configurations.

[0072] Now for reference Figure 4-5 Provided with Figure 1-3 An enlarged cross-sectional view of an engine 10 constructed in a manner similar to that depicted in the illustration of one or more exemplary engines 10. Figure 4-5Walled ducts, manifolds, pipes, or other structures forming a flow path are depicted, the flow path being configured to extract or receive airflow schematically depicted via arrow 91 from compressor section 21 and to provide airflow 91 to turbine section 27. The airflow 91 provided to turbine section 27 can be used to cool blades, impellers, shrouds, or other portions of turbine section 27. In some embodiments, turbine section 27 includes turbine frame 308, which is positioned in a tandem flow arrangement between a first turbine assembly or high-speed turbine 28 and a second turbine assembly or low-speed turbine 30. In further specific embodiments, bearing assembly 200 is included at turbine frame 308. Thus, turbine frame 308 can provide a static mounting or support structure at which bearing assembly 200 is positioned to support the rotation of one or more spools (e.g., low-speed spool 36 or high-speed spool 34). The turbine frame 308 further includes any suitable number of ducts, manifolds, or channels 309, or other structures for allowing at least a portion of the airflow 91 (e.g., further depicted below as airflow 193) to the bearing assembly 200. The airflow to the bearing assembly 200 can provide cooling or buffering air at the bearing assembly 200, such as to dampen vibrations from the spool or to create desired bearing or rotor clearance. In other embodiments, the airflow 91 is provided to a gear assembly 37 located in the fan section 16, to the compressor section 21, to the turbine section 27, or to the exhaust nozzle section 32.

[0073] Engine 10 includes a first duct 110 that extends fluidly from compressor section 21 to turbine section 27. The first duct 110 is configured to connect an airflow 91 from compressor section 21 to a first location 271 at turbine section 27. The first duct 110 forms a flow path separate from the core gas flow path 70. In a particular embodiment, the first duct 110 provides the airflow 91 from compressor section 21 to turbine section 27 while bypassing combustion section 26.

[0074] A first heat exchanger 141 is positioned in thermal communication with an airflow 91 passing through a first duct 110. The first heat exchanger 141 is configured to receive heat or thermal energy from the airflow 91 passing through the first duct 110. Therefore, the first heat exchanger 141 is configured to cool the airflow 91 passing through the first duct 110 before it is 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 or transferring heat or thermal energy to a heat transfer fluid schematically depicted via arrow 221. Specific embodiments of the engine 10 may include a fluid system 220 configured to allow a heat transfer fluid 221 to flow as a lubricant, liquid and / or gaseous fuel, hydraulic fluid, supercritical fluid, refrigerant, or 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. Figure 9 In the specific embodiment depicted (discussed in more detail below), the heat transfer fluid 221 is a liquid fuel supplied to the combustion section 26. However, it should be understood that the heat transfer fluid 221 can be supplied and used in any suitable manner, including but not limited to as a lubricant, anti-icing fluid, fuel, or actuating fluid for bearing systems.

[0075] Still referencing Figure 4-5 The engine 10 includes a second duct 120 extending from the first duct 110 downstream of the first heat exchanger 141 (relative to the airflow 91 from the compressor section 21 to the turbine section 27). The second duct 120 extends fluidly in communication to a second location 272 at the turbine section 27. A flow control device 130 is located at the second duct 120. The flow control device 130 is configured to selectively regulate, alter, modulate, or otherwise change the amount of airflow 91 from the first duct 110 through the second duct 120.

[0076] In various embodiments, the second conduit 120 includes an inlet portion 121 and an outlet portion 122. The inlet portion 121 is fluidly connected 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 connected to the flow control device 130 and a second location 272 at 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 way, it will be understood 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.

[0077] The flow control device 130 may be a valve or any suitable device for adjusting, directing, controlling, or otherwise regulating the amount of fluid flow through a channel or flow path. The flow control device 130 may include an actuated valve or automatic valve driven by electrical power, pneumatic power (e.g., air, or particularly, at least a portion of airflow 91), or a fluid source (e.g., liquid and / or gaseous fuel, hydraulic fluid, lubricant, or combinations thereof). According to the embodiments depicted and described herein, the flow control device 130 may include a ball valve, shuttle valve, or other suitable type of valve or flow regulating device. Thus, the flow control device 130 is configured to regulate the amount of fluid flow through the outlet portion 122 of the second conduit 120, as schematically depicted via arrow 94.

[0078] In a particular embodiment, the engine 10 includes a third duct 123 extending from the flow control device 130 to a third position 273 at the turbine section 27, in fluid communication with both the flow control device 130 and the third position 273. Thus, the flow control device 130 may be a three-way valve configured to selectively alter the amount of airflow 91 from the first duct 110 through the inlet portion 121 of the second duct 120 to one or both of the third duct 123 and the outlet portion 122 of the second duct 120. Therefore, the flow control device 130 may be configured to regulate the amount of airflow 192 through the outlet portion 122 of the second duct 120, as schematically depicted by arrow 194, and further regulate the outflow of at least a portion of the airflow 192 through the third duct 123, as schematically depicted by arrow 195. The third duct 123 may form a bypass passage to further allow selective regulation, control, or modulation of the airflow through the flow control device 130. In a particular embodiment, the third duct 123 allows a portion of the air drawn from the first duct 110 to bypass the outlet portion 122 of the second duct 120 and exit to a third position 273 at the turbine section 27. In some embodiments, the third position 273 allows bypassing the clearance control system 275 (described below) and allows airflow 195 to enter the turbine section 27 at the core gas flow path 70 downstream of the clearance control system 275, or mix with airflow 193 at the turbine frame 308, or vent to the environment (not depicted).

[0079] Still referencing Figure 4-5 As briefly noted above, turbine section 27 includes a clearance control system 275. An exemplary embodiment of the improved clearance control system is described in... Figure 8-16 The structure is depicted as including a housing 300, a manifold assembly 316, and a thermal control ring 314 disposed therein. However, it should be understood that... Figure 4-5The clearance control system 275 depicted may include any suitable structure or component for controlling, regulating, or otherwise modulating the dimension (also referred to as tip clearance) between the rotor blade tip and the surrounding shroud or wall at the turbine section 27. The clearance control system 275 may be an active clearance control (ACC) system configured to dynamically control the tip clearance. Specifically, the ACC system may be configured to desirably regulate the tip clearance based on engine operating conditions, via an airflow 94 received from the second duct 120 and provided to the surrounding shroud at the turbine section 27. The volumetric or mass flow rate of the airflow 94 is adjusted or regulated by a flow control device 130. The amount of airflow 94 regulated to the clearance control system 275 allows the tip clearance to be desirably adjusted under various engine operating conditions and associated temperature variations at the turbine section 27. As temperature and rotor speed vary at the turbine section 27 under various engine operating conditions, the flow control device 130 regulates the amount of airflow 94 provided to the clearance control system 275 to maintain or provide the desired tip clearance. Regarding the landing-takeoff cycle (LTO) of engine 10 and the aircraft, engine operating conditions include start-up, idling, takeoff, climb, cruise, approach, or thrust reverse. However, it should be understood that other engine operating conditions and cycles may apply.

[0080] Still referencing Figure 4-5 The second position 272 at the turbine section 27 is located at the clearance control system 275. Therefore, the second duct 120, or specifically, the outlet portion 122 of the second duct 120, is fluidly coupled to the turbine section 27 to provide airflow 94 to the clearance control system 275, as described herein. In a particular embodiment, the clearance control system 275 is operatively coupled to the first turbine assembly or high-speed turbine 28 at the turbine section 27. Thus, the engine 10 is configured to receive airflow 91 from the compressor section 21 and provide a portion of the airflow 94 (from airflow 91) to the clearance control system 275 at the high-speed turbine 28 via the second duct 120.

[0081] In some other embodiments, the first duct 110 is fluidly connected to a turbine frame 308, which is positioned between a first turbine assembly or high-speed turbine 28 and a second turbine assembly or low-speed turbine 30. The turbine frame 308 may include a plurality of blades 310 arranged circumferentially and positioned between the turbines 28 and 30. A first location 271 at turbine section 27 is located at turbine frame 308. Thus, in these embodiments, the first duct 110 is configured to provide at least a portion of an airflow 91 to turbine frame 308 at the first location 271. In a particular embodiment, schematic arrow 193 depicts a portion of the airflow at the first duct 110 downstream of the junction with the second duct 120. The airflow 193 is provided to turbine frame 308 via the first duct 110. Regarding... Figure 8-16 In a particular embodiment that is further depicted and described, airflow 193 may be provided to housing 300 and through a plurality of blades 310 at turbine frame 308, as schematically depicted via arrow 99.

[0082] refer to Figure 4-5 The turbine frame 308 may include or form one or more channels 309 configured to provide fluid communication of airflow 193 to bearing assembly 200. Airflow 193 may provide buffer fluid for the operation of bearing assembly 200. Buffer fluid may desiccately control or reduce vibration at bearing assembly 200 or at the rotor connecting to the bearing assembly, or allow or generate a desired clearance or vibration response.

[0083] Now for specific reference Figure 5 In a particular embodiment, the engine 10 includes a second heat exchanger 142 in thermal communication with an airflow at a bypass airflow passage 48. The second heat exchanger 142 may be configured as a surface heat exchanger, configured to receive heat or thermal energy from an airflow 194 downstream of a flow control device 130 at the second duct 120. The heat transfer fluid at the second heat exchanger 142 is the airflow through the bypass airflow passage 48 of the engine 10, as schematically 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 position the airflow 194 at the second duct 120 in thermal communication with a bypass airflow 177 at the bypass airflow passage 48. In a particular embodiment, the second heat exchanger 142 is positioned at the outlet portion 122 of the second duct 120 and upstream of a second position 272 at the turbine section 27.

[0084] Generally refer to the reply Figure 4-5Both, in a particular embodiment, the first duct 110 includes an inlet manifold 111 configured to receive an airflow 91 from a circumferential compressor position 211 at the compressor section 21. It should be understood that, although Figure 4-5 The embodiment depicted illustrates a single circumferential compressor position 211, but the inlet manifold can be configured to receive airflows 91 from multiple circumferential compressor positions 211.

[0085] Now for reference Figure 6 Provided according to Figures 1 to 3 A perspective view of an embodiment of one or more of the engines 10. Figure 6 The embodiments provided herein can be configured to relate to... Figure 4-5 The embodiments described herein are essentially similar. Figure 6 In this embodiment, engine 10 may include a plurality of inlet manifolds 111 that are uniformly or asymmetrically spaced in a circumferential direction C around compressor section 21. In various embodiments, the plurality of inlet manifolds 111 may include two (2) or more inlet manifolds. In one embodiment, the plurality of inlet manifolds 111 may include three (3) inlet manifolds. In another embodiment, the plurality of inlet manifolds 111 may include four (4) inlet manifolds, and up to 30 inlet manifolds 111.

[0086] exist Figure 6 In this embodiment, the first duct 110 includes a collector 115 configured to receive an airflow 91 from the inlet manifold 111. In a particular embodiment, a plurality of inlet manifolds 111 are fluidly coupled to a single collector 115 to provide the collected or uniform airflow 91 to a first heat exchanger 141. The collector 115 may provide the airflow 91 to the first heat exchanger 141, as described herein.

[0087] In yet another specific embodiment, the first duct 110 includes an outlet manifold 112 configured to fluidly communicate an airflow 91 from the first heat exchanger 141 to the turbine section 27 at a first turbine location 271 in the turbine section 27. The engine 10 may include a plurality of outlet manifolds 112 spaced uniformly or asymmetrically around the turbine section 27 in a circumferential direction C. 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 duct 120 extends fluidly in communication from one or more of the plurality of outlet manifolds 112 of the first duct 110. Thus, the plurality of outlet manifolds 112 may extend to a plurality of first turbine locations 271 at different circumferential locations in the turbine section 27.

[0088] It should be understood that, although Figure 4-5 The embodiments depicted illustrate a single circumferential first turbine position 271, but the first turbine position 271 may include multiple circumferential first turbine positions 271.

[0089] Figure 4-5 The embodiment of the engine 10 provided herein may include a first duct 110 as a fixed-area flow path from the compressor section 21 to the turbine section 27. In other words, the first duct 110 may include various cross-sectional areas or tapering and expanding flow paths. However, the first duct 110 and the circumferential compressor position 211 may define a fixed or non-connected flow path area. This fixed-area flow path allows an airflow 91 from the compressor section 21 to pass through the first duct 110 at a constant volumetric or mass flow rate relative to a corresponding engine operating condition. In other words, the fixed-area flow path allows the first duct 110 to receive an airflow 91 at a corresponding flow rate relative to a particular engine operating condition. Therefore, the embodiment of the engine 10 provided herein allows a constant airflow 91 to be thermally connected to a heat transfer fluid flow 221 at a 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, may be controlled via a timetable, table, graph, or curve indicating the flow rate relative to engine operating conditions. In one embodiment, the airflow 91 at the first duct 110 may be generally fixed as a ratio or proportion of the total airflow entering the core engine inlet 20 to the compressor section 21. In another embodiment, the airflow 91 at the first duct 110 may be generally fixed as a ratio or proportion of the airflow entering the high-speed compressor 24 from the low-speed compressor 22.

[0090] The engine 10 can specifically include a variable-area flow path at the second duct 120 via the flow control device 130. Thus, the engine 10 can allow a fixed airflow 193 to the turbine frame 308 (such as for the bearing assembly 200) and a variable airflow 194 to the clearance control system 275. The flow control device 130 can adjust, connect, or otherwise regulate the airflow 194 to the clearance control system 275 according to engine operating conditions. Regulation of the airflow 194 via the flow control device 130 can be a function of inlet air velocity (entering the turbine 14 via inlet 20), or inlet air pressure (e.g., corresponding to the altitude of the engine 10 during operation or under one or more of the aforementioned engine operating conditions), or inlet air temperature, or a combination thereof. Alternatively or additionally, regulation of the airflow 194 via the flow control device 130 can be a function of the tip clearance at the turbine section 27, or a predetermined schedule corresponding to wear or degradation at the turbine section 27.

[0091] Some embodiments of engine 10 include a specific pressure range corresponding to the airflow 91 during engine 10 operation, and a specific placement of the circumferential compressor position 211 at a specific axial stage or other location within compressor section 21. In various embodiments, the circumferential compressor position 211 from which it receives the airflow 91 from the core gas flow path 70 corresponds to a compressor position with airflow passing through it at pressure between approximately 20 pounds per square inch (psi) and approximately 60 psi during engine operating conditions, corresponding to approximately 55% and approximately 75% of the operating envelope. In another embodiment, the circumferential compressor position 211 from which it receives the airflow 91 from the core gas flow path 70 may correspond to a compressor position with airflow passing through it at pressure between approximately 30 pounds per square inch (psi) and approximately 50 psi during engine operating conditions such as those described herein.

[0092] Therefore, the embodiments of engine 10 provided herein allow the clearance control system 275 and bearing assembly 200 to operate and receive air from compressor section 21. In some embodiments, the engine 10 provided herein allows the clearance control system 275 to receive airflow 91 from compressor section 21 instead of from bypass airflow passage 48. Furthermore or alternatively, compared to other compressor exhaust systems that may receive high-energy air downstream, behind, or at higher pressure stages of the compressor section, the engine 10 provided herein allows airflow 91 to be received from upstream, in front, or at lower pressure stages of the compressor section. Some of these other compressor exhaust systems may further mix higher-energy air with lower-energy (i.e., lower pressure, lower temperature, or both) air corresponding to the bypass airflow passage. Further or alternatively, the embodiments of engine 10 provided herein allow a constant airflow 91 through the first duct 110 to maintain purging and recirculation margins at turbine frame 308 and bearing assembly 200.

[0093] Now for reference Figures 7A-7B A flowchart outlining the steps of a method 1000 for operating an engine is provided. The steps of method 1000 can be stored as instructions and / or executed as operation via embodiments of the engine 10 and computing system 1210 provided herein. Therefore, method 1000 can be a computer-implemented method in which one or more steps are stored as instructions at memory 1214 at computing system 1210 and / or executed by one or more processors 1212 at computing system 1210. Computing system 1210 can enable operations such as those described herein... Figure 1-6 The described engine embodiments are as follows: Figures 7A-7B The operations outlined in the flowchart and further described herein with respect to method 1000.

[0094] refer to Figures 7A-7B The flowchart in the text, and combined with Figure 1-6 In any one or more embodiments depicted herein, method 1000 includes, at 1010, initiating rotation of one or both of a high-speed spool or a low-speed spool, for example, to generate compressed air for combustion within the combustion section of the core engine. In various embodiments, a motive force, such as a starter motor or a turbo air starter (not shown), initiates rotation of one or both of a high-speed spool 34 or a low-speed spool 36 to generate an initial airflow entering the combustion section 26 through the core gas flow path 70, for mixing with liquid and / or gaseous fuel before ignition to generate combustion gases.

[0095] Method 1000 further includes, at 1020, compressing the airflow via a compressor section. During operation of engine 10, airflow 171 is received at fan section 16. A portion of airflow 171 enters turbine 14 via core engine inlet 20, as schematically depicted via arrow 172. Airflow 172 is pressurized through a continuous row or stage of compressor blades at compressor section 21. In particular, low-speed compressor 22 may include a low-pressure compressor or turbocharger relative to high-speed compressor 24, which includes a high-pressure compressor. In some embodiments, a portion of the airflow 172 compressed by low-speed compressor 22 may be vented or redirected from core gas flow path 70, such as to control stall, surge, or operability at one or both of compressors 22, 24. High-speed compressor 24 receives airflow 172 and further compresses the airflow, such as via... Figure 1-3 Arrow 173 is schematically depicted in the diagram. The continuous stage of compressor blades provides energy to the airflow 173, such as by increasing the pressure and temperature of the airflow 173 before it enters the combustion section 26, as depicted by arrow 174.

[0096] Method 1000 includes, at 1030, extracting a portion of the compressed air flow from the compressor section, as described above. Method 1000 at 1030 may specifically include extracting a portion of the compressed air flow into a first duct and bypassing the combustion section, as provided above with respect to the first duct 110. Method 1000 includes, at 1040, allowing the extracted portion of the compressed air flow to flow through the first duct (e.g., first duct 110) to the turbine section. In a particular embodiment, when allowing the extracted portion of the compressed air flow to flow to the turbine section, the first duct bypasses the combustion section. Regarding Figure 1-6 A portion of the airflow at compressor section 21 is discharged or removed from the core gas flow path 70 and is supplied to the first duct 110, such as via Figure 1-5 Arrow 91 is schematically depicted. In certain embodiments described herein, airflow 91 can be received from compressed airflows 173, 174 from high-speed compressor 24. In still other embodiments, airflow 91 can be received from compressed airflow 172 from low-speed compressor 22.

[0097] It should be understood that the embodiments of engine 10 provided herein advantageously receive a relatively low-pressure and low-temperature airflow from compressor section 21, and can further avoid the structural, complex, actuating device, valve, and associated weight and efficiency losses associated with mixing high-pressure and high-temperature air with low-pressure and low-temperature air from the fan bypass airflow passage. Furthermore, it should be understood that while specific operating conditions and operating envelopes are provided herein, the engine 10 and / or method 1000 provided herein allow one or more steps to be performed under any engine operating condition, including up to 100% of the total power output. However, this document provides specific advantages and benefits regarding engine operation under engine operating conditions that define most of the operating envelope. Thus, the methods and structures provided herein allow for improved efficiency and reduced fuel consumption.

[0098] In various embodiments, 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 a particular embodiment, 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 a particular embodiment, method 1000 includes at 1035 receiving a portion of the compressed air flow from the compressor section, wherein 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 another particular embodiment, method 1000 at 1030 and / or 1035 is performed continuously or constantly relative to discrete engine operating conditions, such as to allow a fixed airflow relative to discrete engine operating conditions.

[0099] In yet another specific embodiment, method 1000 includes operating the engine at 1028, under engine conditions corresponding to approximately 55% and approximately 75% of the operating envelope or approximately 60% and approximately 70% of the operating envelope, as described above. In some embodiments, one or both steps of method 1000 at 1030 and 1035 occur before or simultaneously with method 1000 at 1028. In still other embodiments, method 1000 includes operating the engine at 1029, between approximately 75% and approximately 90% of the engine's total power output (e.g., rated thrust), as described above. In yet another specific embodiment, method 1000 at 1029 includes operating the engine between approximately 80% and approximately 88% of the engine's total power output. In some embodiments, one or more ranges provided herein may define discrete engine operating conditions for continuously or constantly performing method 1000 at 1030 and / or 1035. In some other specific embodiments, method 1000 includes performing the steps at 1028 and 1029 simultaneously.

[0100] Method 1000 may include, at 1050, allowing a heat transfer fluid to flow via a fluid system, thermally communicating with an extraction portion of the compressed air flow, as described above. In a particular embodiment, Figure 4-5 The fluid system 220 depicted is a liquid and / or gaseous fuel system configured to supply a liquid and / or gaseous fuel stream to a compressed air stream 174 to generate combustion gases 175. In this embodiment, the fuel is a heat-transfer fluid 221 in thermal communication with the air stream 91 via a 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 may desiccately affect combustion, fuel-air mixing, swirl, emission generation, vibration, or smoke and particulate generation.

[0101] In some embodiments, method 1000 may further include flowing a plurality of heat transfer fluids in thermal communication with an extraction portion of the compressed air flow via a fluid system. In various embodiments, method 1000 includes providing one or more of a fuel, lubricant, hydraulic fluid, refrigerant, supercritical fluid, or another air flow at the heat transfer fluid in thermal communication with the extracted air flow.

[0102] Method 1000 may further include regulating the heat transfer fluid flow to control the temperature of the extracted air flow (e.g., air flow 91). Regulating the heat transfer fluid flow may include regulating the mass or volumetric flow rate, pressure, or temperature of the heat transfer fluid provided in thermal communication with the extracted air flow.

[0103] As described above, the liquid and / or gaseous fuel stream mixes with compressed air from the compressor section and is ignited to form combustion gas 175. Combustion gas 175 flows from combustion section 26 to turbine section 27, and specifically to high-speed turbine 28 and low-speed turbine 30. As combustion gas 175 expands at turbine section 27, energy is released to drive the rotation of the respective turbines 28, 30, which in turn drives their respective spools 34, 36, compressors 22, 24, and fan blades 40.

[0104] It should be understood that the combustion gases 175 release variable amounts of heat at the turbine section 27 based on engine operating conditions. Therefore, heat release and turbine rotor speed can alter the tip clearance between the turbine rotor blade tips and the surrounding shroud, as further described below. It should be understood that improved aerodynamics and operating efficiency are largely achieved by minimizing the tip clearance. Therefore, a clearance control system is used to adjust the tip clearance based on engine operating conditions to improve engine efficiency and performance.

[0105] Method 1000 may further include, at 1060, selectively directing a portion of the airflow through a second duct (e.g., second duct 120), the second duct (e.g., second duct 120) extending downstream from the first duct (e.g., first duct 110) in a heat exchanger (e.g., first heat exchanger 141). In a particular embodiment, method 1000 includes, at 1062, altering or regulating a portion of the airflow drawn from the first duct (e.g., first duct 110) to the second duct (e.g., second duct 120) downstream of the heat exchanger (e.g., first heat exchanger 141) via a flow control device (e.g., flow control device 130) extending from the first duct. In yet another particular embodiment, method 1000 includes, at 1063, regulating a second portion of the airflow drawn from the first duct to a third duct extending from the flow control device, such as in Figure 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).

[0106] 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.

[0107] 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... Figure 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... Figure 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.

[0108] 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 embodiments described herein are applicable to, for example, Figure 2 When the open rotor configuration described herein is applied, the challenges associated with removing the nacelle and passageways, pipes, or ducts that can be directed through the nacelle to supply air to the heat exchanger, clearance control system, and / or bearing assembly can be overcome. Therefore, method 1000, when applied to an open rotor configuration such as that described herein, can provide a method for operating an open rotor engine, or in particular, a method for operating the clearance control system of an open rotor engine, or more particularly, a method for operating the clearance control system and bearing assembly of an open rotor engine.

[0109] Now for reference Figure 8 Provided according to Figure 1-3 One or more of these can be combined with enlarged cross-sectional views of the turbine section portion of the turbine 14 in various embodiments of this disclosure. For example... Figure 8 As shown, the first turbine assembly is formed by a high-speed turbine 28. The 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 blades 54 (only one shown) axially spaced from the annular array of turbine rotor blades 58 (only one shown) at the high-speed turbine 28. In a particular embodiment, the high-speed turbine 28 further includes a final stage 60, which includes an annular array of stator blades 64 (only one shown) axially spaced from the annular array of turbine rotor blades 68 (only one shown). The turbine rotor blades 58, 68 extend from the high-speed linear shaft 34 (… Figure 1 , Figure 2 Extending radially outward, and connected to the high-speed spool 34 ( Figure 1 , Figure 2 Stator blades 54, 64 and turbine rotor blades 58, 68 at least partially define the core gas flow path 70 for directing gas from combustion section 26 ( Figure 1 , Figure 2 The combustion gases are directed through the high-speed turbine 28.

[0110] like Figure 8As further shown, the high-speed turbine 28 may include one or more shroud assemblies, each shroud assembly forming an annular ring around the rotor blades of the annular array. For example, shroud assembly 72 may form an annular ring around the rotor blades 58 of the annular array of the first stage 50 and the turbine rotor blades 68 of the annular array of the last stage 60. Generally, shroud assembly 72 is radially spaced from the blade tips 76, 78 of each rotor blade 58, 68. The radial or clearance interval CL is defined between the blade tips 76, 78 and the respective inner surfaces of the shroud section 77. Shroud assembly 72 substantially reduces leakage from the core gas flow path 70. Shroud assembly 72 may include multiple walls forming a thermal control ring 314, which helps control thermal growth in the shroud, thereby controlling the radial deflection or clearance interval CL. A clearance control system 275 is used to actively control thermal growth in the shroud assembly. The clearance control system 275 is used to minimize the radial blade tip clearance CL between the outer blade tip and the shroud, particularly during engine cruise operation, as described herein.

[0111] Downstream of the core gas flow path 70 or behind the high-speed turbine 28 is a second turbine assembly formed by the low-speed turbine 30. As previously described herein, the second turbine assembly is rotatably separable from the first turbine assembly, as per the references above. Figure 1 The high-speed turbine 28 and the low-speed turbine 30 are described.

[0112] The housing 300 surrounds the high-speed turbine 28. The housing 300 includes a plurality of blades 310 extending through the core gas flow path 70, behind a first turbine assembly formed by the high-speed turbine 28 and in front of a second turbine assembly formed by the low-speed turbine 30. A shroud assembly 72 is attached to the housing 300 at the outer casing wall 312. The outer casing wall 312 surrounds the shroud assembly 72 and is perpendicular to the centerline axis 12. Figure 1-3 An annular wall extending in the circumferential direction C. The outer casing wall 312 extends in the axial direction A in front of the rotor blades 58 of the first stage 50 of the high-speed turbine 28 (also referred to as the first-stage rotor blades 58) and behind the rotor blades 68 of the second or last stage 60 of the high-speed turbine 28 (also referred to as the second-stage rotor blades 68).

[0113] Multiple blades 310 extend from the housing wall 312. The multiple blades 310 extend into the core gas flow path 70. In some embodiments further described herein, one or more of the multiple blades 310 may be hollow or include conduits or channels allowing fluid flow within the blades. The outer housing wall 312 of the housing 300 extends along the axial direction A from the downstream end or trailing edge of the last stage of the rotor blades 68 to at least the upstream end or leading edge of the multiple blades 310, such as in… Figure 8 The dimension B is depicted in the diagram.

[0114] 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.

[0115] 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).

[0116] 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).

[0117] The integral single structure of the thermal control ring 314 and the housing wall 312, particularly the housing wall extending behind the second or final stage rotor blades 68 of the high-speed turbine 28, allows for improved clearance control, improved thermal control, and improved cooling flow. The structure provided herein allows the thermal control ring 314 to be positioned radially outward and axially aligned with each stage of the high-speed turbine rotor, such as to improve clearance control at each corresponding stage. The structure provided herein further allows for the exclusion of a flange between the high-speed turbine and the intermediate turbine frame, which lies between the high-speed turbine and the downstream low-speed turbine (or, as described herein, the intermediate-speed turbine).

[0118] The embodiments of the monolithic housing provided herein are generally produced by one or more additive manufacturing processes as described above. While additive manufacturing can generally be applied to form a variety of structures or integrate various components, it should be understood that the combination of monolithic structures provided herein can overcome the problems associated with monolithic structures while providing unexpected benefits. In one case, the axially extending housing may be generally susceptible to thermal deformation that could ellipticize the core flow channels, which could adversely affect rotor operation as the rotor may rub within non-concentric flow paths. Thus, a simple integration of a relatively hot housing around a high-speed turbine with a relatively cool housing around the downstream blades of an adjacent low-speed turbine may adversely affect overall engine operation. In another case, this large axially extended mass may require additional cooling flow, which could lead to increased fuel consumption and an overall loss of engine performance.

[0119] The engine embodiments provided herein overcome these problems at least in part by positioning the thermal control ring axially aligned with and radially outward of the corresponding stage of the high-speed turbine blade. Removing the flange between the housing surrounding the high-speed turbine rotor and the downstream blade housing or frame of the high-speed turbine allows the thermal control ring to be advantageously positioned as disclosed herein.

[0120] Other embodiments of the engine provided herein overcome these problems at least in part through improved cooling flow structures, channels, and ducts. In various embodiments, manifold assembly 316 surrounds thermal control ring 314 in the circumferential direction C and the axial direction A. Manifold assembly 316 is configured to provide fluid flow, such as airflow 192 from compressor section 21 (e.g., regarding...) Figure 4-5 (As depicted and described) is provided to the thermal control loop 314.

[0121] Still referencing Figure 8 And now also referencing Figure 9-11 and Figure 14 Further exemplary embodiments are provided. Figure 8 , Figure 9 and Figure 14 The embodiments depicted can be constructed similarly to each other, as further described below. Figure 9-11 Views of fluid flow and openings at various cross-sections of embodiments of the engine 10 at different circumferential locations are provided. Each embodiment can be formed via one or more manufacturing methods known in the art. Figure 14 The embodiments provided herein may include a double-walled structure that can be formed via additive manufacturing. The various embodiments provided herein can be formed into a single, integral structure, such as via additive manufacturing or other suitable manufacturing processes.

[0122] refer to Figure 8-11 and Figure 14 In the various embodiments depicted, the manifold assembly 316 extends along an axial direction A in front of and behind a plurality of axially spaced stages forming a plurality of walls of the thermal control ring 314. In certain embodiments, such as Figure 14 As depicted, the manifold assembly 316 extends rearward along the axial direction A of the plurality of blades 310. In various embodiments, such as in Figure 8 In exemplary embodiments, the manifold assembly 316, the housing wall 312, and the plurality of walls of the thermal control ring 314 forming the housing 300 are a single, integral, monolithic structure, as described herein. In certain embodiments, such as in Figure 8 In an exemplary embodiment, the manifold assembly 316 includes a plurality of concentric walls integrally formed and surrounding the housing wall 312. In some embodiments, the manifold assembly 316 includes an inner manifold 1316 radially inside and concentric with the outer manifold 2316. In still other embodiments, the inner manifold 1316 is a double-walled structure concentric with the outer manifold 2316.

[0123] Special Reference Figure 9-10 Some embodiments of housing 300 include a corrugated feature 399. The corrugated feature 399 includes a shape defining a ridge or recess configured to mitigate the formation of thermal expansion stress at housing 300. In some embodiments, the corrugated feature 399 is formed at manifold assembly 316. In yet another specific embodiment, the corrugated feature 399 may be formed at inner manifold 1316 or outer manifold 2316. The corrugated feature 399 may allow for a single, integral formation of manifold assembly 316 with housing wall 312, as described in the various embodiments herein.

[0124] Now for a brief reference Figure 15The manifold assembly 316 includes a plurality of openings 318 that surround a plurality of walls forming a thermal control ring 314 at the housing 300. The plurality of openings 318 allow a fluid flow schematically depicted via arrow 91 to be in thermal communication with the thermal control ring 314 for a desired heat transfer effect. In various embodiments, the plurality of openings 318 includes an inlet opening 3181 configured to allow an airflow 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 may further include an outlet opening 3182 configured to allow at least a portion of the airflow 91 schematically depicted via airflow 92 to exit the first cavity 1321 and enter an inner wall conduit 1326, as further described below.

[0125] An inlet opening wall 381 extends between the outer portion 346 and the inner portion 347 of a 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 fluidly separated from the inner wall duct 1326. The inlet opening 3181 and the inlet opening wall 381 allow airflow 91 to pass through the duct 1324 surrounding the inner manifold 1316 and enter the air chamber 383 formed between adjacent thermal control rings 314. Specifically, 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 airflow 91 to enter the air chamber 383 while being fluidly isolated from airflow 92 by the inner wall duct 1326.

[0126] Special Reference Figure 9-10 As discussed, manifold assembly 316 includes an inner manifold 1316 surrounding a thermal control ring 314 in both the circumferential direction C and the axial direction A. The depicted manifold assembly 316 further includes an outer manifold 2316 surrounding the inner manifold 1316, as discussed above. A channel wall 1318 extends from the inner manifold 1316 to the outer manifold 2316 to form a channel 1320 within the channel wall 1318.

[0127] In some embodiments, such as Figure 8 As depicted, the outer manifold 2316 of the manifold assembly 316 extends along the axial direction A at or behind the plurality of blades 310. The outer manifold 2316 is further connected to the outer housing wall 312 at or behind the plurality of blades 310. In some other embodiments, such as Figure 9-11As depicted, the internal manifold 1316 extends forward (terminating in front of the plurality of blades 310) along the axial direction A of the plurality of blades 310. The internal manifold 1316 also extends backward along the axial direction A of the plurality of walls forming the thermal control ring 314. Thus, the internal manifold 1316 is connected to the housing wall 312 in front of the plurality of blades 310 and behind the thermal control ring 314.

[0128] The above reference Figure 15 (also Figure 9-11 The first cavity 1321 (described in the diagram) is formed between the inner manifold 1316 and the outer casing wall 312. A thermal control ring 314 is located within the first cavity 1321 between the inner manifold 1316 and the outer casing wall 312, surrounded by the inner manifold 1316. A channel 1320 allows fluid communication with the first cavity 1321 between the inner manifold 1316 and the outer casing wall 312. The channel 1320 further allows an airflow 91 to enter thermal communication with the thermal control ring 314.

[0129] In various embodiments, the conduit 1324 briefly mentioned above is formed between the external manifold 2316 and the internal manifold 1316. The conduit 1324 is in fluid communication with the first lumen 1321 and is fluidly separated from the channel 1320 by the channel wall 1318. In a particular embodiment, the channel wall 1318 extends from the external manifold 2316 to the internal manifold 1316 through the conduit 1324.

[0130] Special Reference Figure 9-11 And further regarding Figure 14 The duct 1324 extends further in fluid communication through one or more of the plurality of blades 310. Figure 10 and Figure 14 Specifically, an airflow 91 that enters thermal and fluid communication with the thermal control ring 314 in the first cavity 1321 is depicted. Figure 10 Specifically, an airflow 91 is depicted that enters thermal and fluid communication with the thermal control ring 314 in the first cavity 1321. In various embodiments, the first cavity 1321 is configured to direct the fluid flow directly to the thermal contact portion of the thermal control ring, for example, in the vertical direction. Figure 11 and Figure 14 Specifically, an airflow 92 (as discussed below) is depicted exiting the first cavity 1321 via duct 1324 and subsequently flowing continuously through one or more of the plurality of impeller blades 310. In some embodiments, a thermal control ring 314 is formed together with the housing wall 312 to desirously improve clearance control. In one embodiment, such as Figure 13B As depicted, the thermal control ring 314 includes an outer surface that extends as a ridge, a groove, or at an acute angle or zigzag angle (see the more detailed description below).

[0131] Brief Special Reference Figure 14 And in Figure 15 As further depicted in the detailed perspective view, in some embodiments, the inner manifold 1316 is a double-walled structure forming an inner wall duct 1326 between the double-walled structures of the inner manifold 1316. The inner wall duct 1326 can extend fluidly in communication to a second cavity 1322, which is formed between the outer casing wall 312 and the outer wall 170 of the core gas flow path 70. In these embodiments, the single integral housing 300, or further, the integration of embodiments of the manifold assembly 316, allows for separate flows into multiple impellers 310. Specifically, airflow 91 originates from, for example, from... Figure 1-6 The compressor section depicted and described enters duct 1324. A portion of the airflow 91, depicted by arrow 92, flows into the first chamber 1321 and then into the inner wall duct 1326 formed at the double-walled structure. The airflow 92 then flows into one or more of the plurality of impeller blades 310. Furthermore, another portion of the airflow 91, depicted by arrow 99, remains in duct 1324 and flows into one or more of the plurality of impeller blades 310. In some embodiments, flows 92, 99 are isolated from each other or fluidly separated until they mix at the plurality of impeller blades 310. In other embodiments, flows 92, 99 are maintained fluidly separated and provided to separate the respective impeller blades 310 or to separate the ducts within each impeller blade 310. Embodiments of the housing 300 and manifold assembly 316, such as by providing secondary use of the fluid flow after thermal communication with the thermal control ring 314, rather than outputting the flow to the atmosphere or the under-shroud area of ​​the engine, allow for improved thermal efficiency and improved overall engine efficiency.

[0132] In some embodiments, the outer wall 170 of the core gas flow path 70 forms an outer shield section 77 of the shield assembly 72. The outer shield section 77 is exposed to the core gas flow path 70 and may include a thermal barrier coating or material configured to withstand heat from the combustion gases. The outer shield section 77 may be further configured to rub against at least partially one or more stages of blades at the core gas flow path 70.

[0133] Still referencing Figure 14 And further in Figure 15 Provided Figure 15 Side view of housing 300 Figure 16 and provided Figure 16 A close-up view of section A in the middle. Figure 17As depicted, the inner manifold 1316 includes a chamber wall 1319 extending from the inner manifold 1316 and surrounding the thermal control ring 314. In some embodiments, the chamber wall 1319 extends radially inward from an inner portion 347 of the inner manifold 1316. The chamber wall 1319 may be formed integrally or as a single, monolithic structure with the inner manifold 1316, which includes an outer portion 346 and an inner portion 347. A first cavity 1321 is formed between the outer surface 1325 of the thermal control ring 314 and the chamber wall 1319.

[0134] Special Reference Figure 16 and 17 The thermal control ring 314 includes a wall or body 332 extending outward from the housing wall 312 (e.g., outward in the radial direction R). In various embodiments, such as those described above with respect to multiple thermal control rings 314, the body 332 extends in the circumferential direction C ( Figure 1-3 It extends in a basically circular pattern.

[0135] More specific reference Figure 17 The body 332 forms an internal flow path 330 to allow fluid flow through the thermal control ring 314. The fluid flow through the body 332 allows for desired control, alteration, or modulation of the temperature or thermal gradient at the thermal control ring 314 by varying the temperature or flow rate of the fluid flow through the flow path 330 at the body 332. Furthermore, the fluid flow through the body 332 can allow one or more structures attached to or integrally formed into the thermal control ring 314, such as the outer casing wall 312 or the shroud assembly 72, to move at least in part based on thermal changes provided by the fluid flow, such as to desired control the gap spacing CL between the rotor blades 58, 68 and the shroud assembly 72. Figure 8 ).

[0136] Still referencing Figure 17 The illustrated exemplary housing 300 further includes a plurality of pins 334 extending from the outer housing wall 312 to the body 332 along a radial direction R of the engine 10 to which the housing 300 is attached. Also briefly referenced... Figure 18 The top-down view of multiple pins 334 depicts each pin 334. (See attached image.) Figure 17 and 18 As shown, each pin 334 is spaced apart from each other along the axial direction A of the engine 10 of the connecting housing 300 and along the circumferential direction C of the engine 10 of the connecting housing 300. Figure 18 In this way, adjacent pins 334 define a gap 336 between them.

[0137] Special Reference Figure 17The radially extending flow path 330 through the body 332 further fluidly extends into the gaps or voids 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 to each other 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. Based on the above regarding... Figure 1-15 In any one or more embodiments depicted and described, an airflow schematically depicted via arrow 91 is received and provided in fluid communication with thermal control ring 314.

[0138] During operation, airflow 91 passes through gap 336 and through multiple pins 334 to enter flow path 330 within body 332. During operation, airflow 91 advances radially through body 332 and exits body 332 through outlet opening 338 at flow path 330. Outlet opening 338 is formed by body 332 away from gap 336 to allow fluid communication from flow path 330 to inner wall duct 1326 formed within the double-walled structure of inner manifold 1316. According to... Figure 1-15 In any one or more embodiments depicted and described, fluid flow from the thermal control ring 314, schematically depicted via arrow 92, can flow through the inner wall conduit 1326.

[0139] Still referencing Figure 17 In various embodiments, the seal 1323 is positioned to contact the outer surface 1325 of the thermal control ring 314 and the chamber wall 1319. Alternatively, the seal 1323 may 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 prevents fluid flow through the first cavity 1321. In certain embodiments, the seal 1323 may be formed as a structural member configured to provide structural support to the inner manifold 1316 and / or the thermal control ring 314. The seal 1323 may further support the body 332 relative to a plurality of pins 334. In some embodiments, the seal 1323 is a brazed, welded, or other component that attaches the chamber wall 1319 to the thermal control ring 314 at the first cavity 1321. It should be understood that the seal 1323 and the chamber wall 1319 may each extend substantially in the same direction as the thermal control ring 314 as an integral annular component or as multiple arcuate sections positioned in a circumferential arrangement.

[0140] In certain embodiments, the body 332 of the outer casing wall 312, the plurality of pins 334, and the thermal control ring 314 is a single integral structure, which may be formed by additive manufacturing or other suitable manufacturing processes. In other specific embodiments, the inner portion 347, the outer portion 346, and the chamber wall 1319 are formed together as a single integral structure of the inner manifold 1316. In some embodiments, the thermal control ring 314 and the outer casing wall 312 are a single structure separate from the inner manifold 1316. In still some embodiments, the single structure is formed by additive manufacturing.

[0141] Now for reference Figure 19 An exemplary embodiment describing the operation of engine 10 is provided. Figure 19 The embodiments provided are configured to be substantially similar to those regarding Figure 16 The embodiments described herein are illustrated. The operation of the system provided herein can be substantially based on the embodiments described with respect to engine 10, such as those concerning... Figure 1-6 and Figures 7A-7B What is depicted and described. In Figure 19 In this configuration, airflow 91 is received at a second position 272 (such as through an opening provided by the outer manifold 2316). Airflow 91 is received in a duct 1324 formed between the outer manifold 2316 and the inner manifold 1316. Airflow 91 is directed into the air chamber 383 via an inlet opening 1381 formed through the inner manifold 1316. Airflow 91 is directed through a plurality of pins 334 and through a flow path 330 (see...). Figure 17 ), enter the inner wall catheter 1326 (see Figure 17 ).

[0142] In one embodiment, such as Figure 19 As depicted, airflow 92 can exit from the inner wall duct 1326 through opening 1380 to the outside of housing 300 or engine 10, as illustrated by arrow 93. Airflow 93 can allow heat or thermal energy from thermal control ring 314 to escape to atmospheric conditions, or to the area under the housing or shroud.

[0143] Now for reference Figure 20 It provides a perspective view of a portion of engine 10. Figure 20 The embodiments provided are configured to be substantially similar to those regarding Figure 16-19 The described embodiments. In particular, Figure 20Multiple discrete flow paths 330 extending through the thermal control ring 314 in an adjacent circumferential arrangement are depicted. Multiple outlet openings 3182 are formed through internal branches 347 of the internal manifold 1316 corresponding to the multiple flow paths 330 and outlet openings 338 at the thermal control ring 314. Therefore, the engine 10 can form multiple flow paths 330 and outlet openings 338 adjacently arranged in the circumferential direction C at the thermal control ring 314, corresponding to the multiple outlet openings 3182 formed through the internal branches 347 of the internal manifold 1316. This arrangement allows airflow 92 to exit from inside and outside the thermal control ring 314 into the inner wall duct 1326.

[0144] Now for reference Figure 21 Provided Figure 20 A side cross-sectional view of the embodiment provided. Figure 21 The embodiment further depicts an inner wall duct 1326 in fluid communication with a second cavity 1322 located at turbine frame 308. An opening 3112 is formed through turbine frame 308 to allow airflow 92 to exit and is in thermal communication with turbine frame 308.

[0145] Now, let's briefly refer to the previous section. Figure 12 and Figures 13A-13D Additional aspects of this disclosure are described. Figure 12 A partial circumferential view of an embodiment of the manifold assembly 316 is provided. Furthermore, Figures 13A-13D Provided Figure 12 Cross-sectional view of the embodiment depicted in the figure (for Figures 13A-13D The label of each in Figure 12 (As indicated in the text). As previously described, various embodiments of the manifold assembly 316 are formed via one or more additive manufacturing processes. See in particular... Figure 12 and Figure 13C A close-up view shows that, in various embodiments, member 3316 extends into the inner manifold 1316 and the outer manifold 2316. Member 3316 extends from the inner manifold 1316 to the outer manifold 2316 at an acute angle (e.g., V-, Z-, or other angled cross-section). In various embodiments, member 3316 extends along a first direction schematically depicted via arrow 95 and along a second direction opposite to the first direction schematically depicted via arrow 96.

[0146] The improved turbine housing 300, turbine section 27, and engine 10 embodiments provided herein allow for improved clearance control, cooling fluid distribution, reduced weight, and improved engine efficiency. The embodiments of engine 10, housing 300, and manifold assembly 316 provided herein include a single, integral structure, such as a housing extending over multiple stages of a high-speed turbine, or further including an inter-turbine frame, or further including all or part of a manifold, such as that which can be formed by additive manufacturing processes that have been impossible or infeasible to date. The embodiments depicted and described herein allow for improved and advantageous positioning of the thermal control ring 314, through which flow paths 330 and multiple pins 334 are located, for improved clearance control response, improved formation and positioning of openings, channels, and conduits to allow for more efficient utilization and movement of heat transfer fluids, and weight reduction, such as by eliminating flanges and sub-assemblies as integral parts. Specific combinations of these features allow for improved heat transfer properties and reduced thermal gradients. Improved heat transfer properties, in particular, include reduced heat transfer coefficients at certain features, such as the multiple walls, body, pins, and / or flow paths forming the thermal control ring 314. This improvement can mitigate or eliminate undesirable or excessive deformation, ellipticization, bending, or other changes in the geometry of the housing 300 that may adversely affect deflection or cause undesirable contact with the turbine rotor blades 58 at the high-speed turbine 28.

[0147] The embodiments of engine 10 and housing 300 provided herein include an integral single housing for a high-speed turbine 28, together with a turbine center frame or intermediate turbine frame 308, formed by housing walls 312 and a plurality of blades 310, and positioned downstream of the core gas flow path 70 of the high-speed turbine 28 and upstream of the core gas flow path 70 of the low-pressure or intermediate-pressure turbine, as depicted at turbine 30. The embodiments provided herein further include, for example, an integral single clearance control manifold configured to provide heat transfer fluid to a thermal control ring. The integral single structure may further allow for improved positioning of the thermal control ring relative to the turbine rotor, such as providing improved clearance control across the turbine rotor assembly.

[0148] It should be understood that, regarding Figure 1-6 The conduits 110, 120, 123, flow control device 130, or heat exchangers 141, 142 depicted and described may be provided to the housing 300, manifold assembly 316, and related components. Figure 8-21Other structures depicted and described herein. However, various embodiments of the engine 10 provided herein may include one or more of the ducts 110, 120, 123, flow control device 130, or heat exchangers 141, 142 that provide airflow to any suitable clearance control system, turbine section, or bearing assembly. Such a structure, when combined with any suitable clearance control system, turbine section, or bearing assembly, may provide one or more of the advantages and benefits described herein. Alternatively, various embodiments of the engine 10 provided herein may include one or more of the housing 300 or manifold assembly 316 that receive airflow from any suitable duct, passage, flow path, pipe, or other structure. Such a structure, when combined with any suitable duct or heat exchanger, may provide one or more of the advantages and benefits described herein. The benefits and advantages described with respect to ducts, flow control devices, heat exchangers, housings, or manifolds, when combined together, may combine the benefits and advantages described herein.

[0149] The embodiments of conduits 110, 120, 123 and heat exchangers 141, 142 provided herein can be formed at least in part by one or more additive manufacturing processes such as those described herein. For example, a first heat exchanger 141 can be integrally formed with a first conduit 110, or a second heat exchanger 142 can be integrally formed with a second conduit 120 or a portion thereof. In another example, all or a portion of a first conduit 110, including one or more inlet manifolds 111, outlet manifolds 112, or collectors 115, can be integrally formed as a single component. In yet another example, all or a portion of a second conduit 120, including one or more inlet portions 121 or outlet portions 122, can be formed as a single component. Furthermore, certain combinations of portions of the first conduit 110, the second conduit 120, and the third conduit 123 can be integrally formed with each other. For example, an outlet manifold 112 can be formed as a single component with an inlet portion 121. In another example, the housing surrounding the compressor section 21 can be integrally formed with the inlet manifold 111. Collector 115 may be integrally formed with first heat exchanger 141. Second heat exchanger 142 may be integrally formed with outlet portion 122.

[0150] This written description uses examples to disclose preferred embodiments, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system, and any methods of combining them. The patentable scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. These other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0151] Further aspects of this disclosure are provided by the subject matter of the following provisions:

[0152] A method of operating a gas turbine engine, the gas turbine engine having a compressor section and a turbine section arranged in series, a first duct in fluid communication with the compressor section and the turbine section, a first heat exchanger positioned in thermal communication with an airflow through the first duct, a second duct in fluid communication with the first duct at a downstream position of the first heat exchanger and at a second position in the turbine section, and a flow control device positioned in fluid communication with the second duct, the method comprising: extracting an airflow from the compressor section into the first duct; allowing the extracted airflow to flow through the first duct to a first position in the turbine section, wherein the second duct is in fluid communication with the turbine section at the second position; allowing a heat transfer fluid to flow to the first heat exchanger, the heat transfer fluid being in thermal communication with the extracted airflow through the first duct via the first heat exchanger; and regulating a portion of the airflow extracted from the first duct to the second duct downstream of the first heat exchanger via the flow control device.

[0153] The method described according to one or more of these provisions further includes operating the engine between approximately 75% and approximately 90% of the engine's total power output.

[0154] The method according to one or more of these provisions, wherein extracting the airflow from the compressor section includes extracting the airflow from the compressor section at a compressor location having an airflow pressure between approximately 20 psi and approximately 60 psi, while operating the engine between approximately 75% and approximately 90% of the total power output.

[0155] The method according to one or more of these provisions further includes: operating the engine in an operating condition corresponding to approximately 55% and approximately 75% of the operating envelope; and wherein extracting the airflow from the compressor section includes, during the operating condition corresponding to approximately 55% and approximately 75% of the operating envelope, extracting the airflow from the compressor section at a compressor location having an airflow pressure between approximately 20 psi and approximately 60 psi.

[0156] According to one or more of these provisions, the compressor section includes a low-speed compressor and a high-speed compressor, and the turbine section includes a low-speed turbine, a high-speed turbine and a turbine frame, the turbine frame being positioned between the low-speed turbine and the high-speed turbine, and the first position at the turbine section being at the turbine frame, and the first duct extending fluidly from the high-speed compressor to the turbine frame.

[0157] According to one or more of these provisions, the first duct is configured as a fixed-area flow path from the compressor section to the turbine section, and the flow control device defines a variable-area flow path at the second duct, and the extraction of airflow through the first duct to the first position at the turbine section is a continuous flow of the engine's operating conditions, and the regulation of the portion of the airflow extracted from the first duct to the second duct includes providing a variable flow to the second position at the turbine section.

[0158] According to one or more of these provisions, the gas turbine engine further includes: a gap control system positioned at the second location in the turbine section; and the method further includes: selectively changing the tip gap at the gap control system based on a portion of the airflow extracted from the first duct to the second duct.

[0159] The method according to one or more of these provisions, wherein the gas turbine engine includes: a third duct extending from the flow control device and in fluid communication with a third location at the turbine section; wherein the method further includes: regulating a second portion of the airflow drawn from the first duct to the third duct extending from the flow control device via the flow control device.

[0160] According to one or more of these provisions, the gas turbine engine further comprises: a fan section, wherein a bypass airflow passage is formed downstream of the fan section and surrounds a housing surrounding the compressor section and the turbine section; and a second heat exchanger, the second heat exchanger being positioned downstream of the flow control device at the second duct and upstream of the second location at the turbine section, wherein the second heat exchanger allows a bypass airflow from the bypass airflow passage to be thermally connected to a portion of the airflow extracted to the second duct; wherein the method further comprises: thermally connecting the bypass airflow to the portion of the airflow extracted to the second duct via the second heat exchanger.

[0161] A gas turbine engine includes: a compressor section and a turbine section arranged in series; a first duct extending from the compressor section to the turbine section, the first duct being in fluid communication with both the compressor section and the turbine section to communicate an airflow from the compressor section to a first location at the turbine section; a first heat exchanger positioned in thermal communication with the airflow through the first duct; a second duct in fluid communication with the first duct at a location downstream of the first heat exchanger and in fluid communication with a second location at the turbine section; and a flow control device positioned in fluid communication with the second duct for selectively changing the amount of the airflow from the first duct through the second duct.

[0162] According to one or more of these provisions, the gas turbine engine wherein the second duct includes an inlet portion and an outlet portion, wherein the inlet portion is fluidly connected to the first duct and the flow control device, and wherein the outlet portion is fluidly connected to the flow control device and the second location of the turbine section.

[0163] The gas turbine engine according to one or more of these provisions includes: a second heat exchanger in thermal communication with an airflow through the outlet portion of the second duct, and upstream of the second location in the turbine section.

[0164] According to one or more of these provisions, the gas turbine engine includes: a fan section, wherein a bypass airflow passage is formed downstream of the fan section and surrounds a housing surrounding the compressor section and the turbine section, wherein the fan section provides a bypass airflow to the bypass airflow passage during operation of the gas turbine engine, and wherein a second heat exchanger provides thermal communication between the bypass airflow at the bypass airflow passage and the airflow at the outlet portion of the second duct.

[0165] The gas turbine engine according to one or more of these provisions includes: a third duct that extends fluidly from the flow control device to a third location at the turbine section.

[0166] According to one or more of these provisions, the gas turbine engine wherein the flow control device is a three-way valve configured to selectively change the amount of airflow from the first duct through the inlet portion of the second duct, and wherein the flow control device is configured to allow at least a portion of the airflow to exit to the third duct, the outlet portion of the second duct, or both.

[0167] According to one or more of these provisions, the gas turbine engine, the turbine section includes: a first turbine assembly surrounded by an outer wall forming a gas flow path, wherein the first turbine assembly includes an outer housing wall surrounding the outer wall, wherein the outer wall and the outer housing wall together form a cavity, and wherein the third position at the turbine section is located at the cavity.

[0168] According to one or more of these provisions, the gas turbine engine includes: a first turbine assembly; a clearance control system, wherein the second position at the turbine section is located at the clearance control system and wherein the clearance control system is operatively coupled to the first turbine assembly; and a turbine frame, the turbine frame being positioned downstream of the first turbine assembly in a tandem flow arrangement, wherein the first position at the turbine section is located at the turbine frame.

[0169] The gas turbine engine according to one or more of these provisions includes: a fluid system configured to provide a heat transfer fluid flow that is in thermal communication with the airflow via the first heat exchanger.

[0170] According to one or more of these provisions, the gas turbine engine, wherein the first duct comprises: a plurality of inlet manifolds configured to receive the airflow from a plurality of circumferential compressor locations in the compressor section; and a collector configured to receive the airflow from the plurality of inlet manifolds, wherein the collector is fluidly coupled to the first heat exchanger to provide the airflow in thermal communication with the first heat exchanger.

[0171] An airflow delivery system for a gas turbine engine including a compressor section and a turbine section arranged in series, the airflow delivery system comprising: a first duct configured to extend from the compressor section to the turbine section and in flow communication with both the compressor section and the turbine section to, when installed in the gas turbine engine, communicate an airflow from the compressor section to a first location at the turbine section; a first heat exchanger positioned in thermal communication with the airflow at the first duct; a second duct extending downstream of the first heat exchanger from the first duct, wherein the second duct is configured to extend in flow communication to a second location at the turbine section; and a flow control device positioned at the second duct, wherein the flow control device is configured to selectively change the amount of airflow passing from the first duct through the second duct.

Claims

1. A method of operating a gas turbine engine, the gas turbine engine having a compressor section and a turbine section arranged in series, a first duct in fluid communication with the compressor section and the turbine section, a first heat exchanger positioned in thermal communication with an airflow through the first duct, a second duct downstream of the first heat exchanger in fluid communication with the first duct and with the turbine section, a flow control device positioned in fluid communication with the second duct, and the second heat exchanger configured as a surface heat exchanger, characterized in that, The method includes: Airflow from the compressor section is drawn into the first duct; The extracted airflow is made to flow through the first duct to a first position in the turbine section, wherein the second duct is in fluid communication with the turbine section at a second position, wherein the second heat exchanger is located downstream of the flow control device at the second duct and upstream of the second position in the turbine section; A heat transfer fluid is directed to the first heat exchanger, the heat transfer fluid being in thermal communication with the extracted airflow passing through the first conduit, wherein the heat transfer fluid is liquid fuel supplied to the combustion zone. A portion of the airflow drawn from the first duct to the second duct downstream of the first heat exchanger is regulated via the flow control device; and The bypass airflow from the bypass airflow channel is thermally connected to a portion of the airflow extracted into the second duct via the second heat exchanger.

2. The method according to claim 1, characterized in that, Further includes: The engine is operated between approximately 75% and approximately 90% of its total power output.

3. The method according to claim 2, characterized in that, Extracting the airflow from the compressor section includes extracting the airflow from the compressor section at a compressor location having an airflow pressure between approximately 20 pounds per square inch and approximately 60 pounds per square inch, while operating the engine between approximately 75% and approximately 90% of the total power output.

4. The method according to claim 1, characterized in that, Further includes: The engine is operated under operating conditions corresponding to approximately 55% and approximately 75% of the operating envelope; Extracting the airflow from the compressor section includes extracting the airflow from the compressor section at a compressor location with an airflow pressure between approximately 20 psi and approximately 60 psi during the operating conditions corresponding to approximately 55% and approximately 75% of the operating envelope.

5. The method according to claim 1, characterized in that, The compressor section includes a low-speed compressor and a high-speed compressor, and the turbine section includes a low-speed turbine, a high-speed turbine and a turbine frame, the turbine frame being positioned between the low-speed turbine and the high-speed turbine, and the first position at the turbine section being at the turbine frame, and the first duct extending fluidly from the high-speed compressor to the turbine frame.

6. The method according to claim 1, characterized in that, The first duct is configured as a fixed-area flow path from the compressor section to the turbine section, and the flow control device defines a variable-area flow path at the second duct, and the extraction of airflow through the first duct to the first position at the turbine section is a continuous flow of the engine's operating conditions, and the regulation of the portion of the airflow extracted from the first duct to the second duct includes providing a variable flow to the second position at the turbine section.

7. The method according to claim 1, characterized in that, The gas turbine engine further includes: Clearance control system, wherein the clearance control system is located at the second position in the turbine section; and The method further includes: Based on a portion of the airflow extracted from the first conduit to the second conduit, the tip gap is selectively changed at the gap control system.

8. The method according to claim 1, characterized in that, The gas turbine engine mentioned above includes: A third conduit extends from the flow control device and is in fluid communication with a third location at the turbine section; The method further includes: The second portion of the airflow drawn from the first conduit to the third conduit extending from the flow control device is regulated via the flow control device.

9. The method according to claim 1, characterized in that, The gas turbine engine further includes: A fan section, wherein a bypass airflow passage is formed downstream of the fan section and surrounds the outer casing surrounding the compressor section and the turbine section.

10. A gas turbine engine, characterized in that, The gas turbine engine includes: The compressor section and turbine section are arranged in series flow. A first duct extends from the compressor section to the turbine section, and is in fluid communication with both the compressor section and the turbine section to connect an airflow from the compressor section to a first location in the turbine section. A first heat exchanger is positioned in thermal communication with the airflow passing through the first duct. The second conduit is in fluid communication with the first conduit at a location downstream of the first heat exchanger and at a second location in the turbine section; A flow control device, positioned in flow communication with the second conduit, for selectively changing the amount of airflow passing from the first conduit through the second conduit; A second heat exchanger, configured as a surface heat exchanger, is positioned downstream of the flow control device at the second duct and upstream of the second location at the turbine section, wherein the second heat exchanger allows a bypass airflow from a bypass airflow passage to be in thermal communication with a portion of the airflow from the second duct; and A fluid system configured to provide a heat transfer fluid flow in thermal communication with the airflow via the first heat exchanger and the first duct, wherein the heat transfer fluid is liquid fuel provided to the combustion zone.

11. The gas turbine engine according to claim 10, characterized in that, The second conduit includes an inlet portion and an outlet portion, wherein the inlet portion is fluidly connected to the first conduit and the flow control device, and wherein the outlet portion is fluidly connected to the flow control device and the second location of the turbine section.

12. The gas turbine engine according to claim 10, characterized in that, The gas turbine engine includes: A fan section, wherein a bypass airflow passage is formed downstream of the fan section and surrounds the housing surrounding the compressor section and the turbine section, wherein the fan section provides bypass airflow to the bypass airflow passage during operation of the gas turbine engine.

13. The gas turbine engine according to claim 10, characterized in that, The gas turbine engine includes: A third conduit extends fluidly from the flow control device to a third location in the turbine section.

14. The gas turbine engine according to claim 13, characterized in that, The flow control device is a three-way valve configured to selectively change the amount of airflow from the first conduit through the inlet portion of the second conduit, and the flow control device is configured to allow at least a portion of the airflow to exit to the third conduit, the outlet portion of the second conduit, or both.

15. The gas turbine engine according to claim 13, characterized in that, The turbine section includes: A first turbine assembly, the first turbine assembly being surrounded by an outer wall forming a gas flow path, and wherein the first turbine assembly includes an outer housing wall surrounding the outer wall, wherein the outer wall and the outer housing wall together form a cavity, and wherein the third position at the turbine section is located in the cavity.

16. The gas turbine engine according to claim 10, characterized in that, The turbine section includes: First turbine assembly; A clearance control system, wherein the second position at the turbine section is located at the clearance control system, and wherein the clearance control system is operatively coupled to the first turbine assembly; and A turbine frame, the turbine frame being positioned downstream of the first turbine assembly in a tandem flow arrangement, wherein the first position at the turbine section is located at the turbine frame.

17. The gas turbine engine according to claim 10, characterized in that, The first catheter includes: Multiple inlet manifolds, the multiple inlet manifolds being configured to receive the airflow from multiple circumferential compressor positions at the compressor section; and A collector configured to receive the airflow from the plurality of inlet manifolds, wherein the collector is fluidly coupled to the first heat exchanger to provide the airflow in thermal communication with the first heat exchanger.

18. An airflow delivery system for a gas turbine engine, the gas turbine engine comprising a compressor section and a turbine section arranged in series, characterized in that, The airflow delivery system includes: A first duct, configured to extend from the compressor section to the turbine section and flow in communication with the compressor section and the turbine section, to, when installed in the gas turbine engine, connect an airflow from the compressor section to a first location at the turbine section; A first heat exchanger is positioned in thermal communication with the airflow at the first duct. A second conduit extends downstream of the first conduit from the first conduit, wherein the second conduit is configured to extend in fluid communication to a second location at the turbine section; A flow control device, located at the second duct, wherein the flow control device is configured to selectively change the amount of airflow from the first duct through the second duct; A second heat exchanger, configured as a surface heat exchanger, is positioned downstream of the flow control device at the second duct and upstream of the second location at the turbine section, wherein the second heat exchanger allows a bypass airflow from a bypass airflow passage to be in thermal communication with a portion of the airflow from the second duct; and A fluid system configured to provide a heat transfer fluid flow in thermal communication with the airflow via the first heat exchanger and the first duct, wherein the heat transfer fluid is liquid fuel provided to the combustion zone.

Images