CN118815592A - Active clearance control assembly - Google Patents

Abstract

A gas turbine engine includes a fan section, an engine inlet, and a fan duct splitter in serial flow order. The fan duct splitter splits an airflow entering the engine inlet from the fan section to the fan duct and the core duct. The core duct includes a compressor section, a combustion section, and a turbine section in serial flow order. The duct assembly is coupled to the fan duct to extract a portion of the fan duct airflow through the fan duct and to deliver the portion of the fan duct airflow to the active clearance control mechanism of the turbine section.

Description

Active clearance control assembly

PRIORITY INFORMATION

The present application claims priority from the polish patent of application number p.444447 filed at month 4 of 2023, 18.

Technical Field

The present subject matter relates generally to components of gas turbine engines, or more particularly, to an active clearance control assembly.

Background

Gas turbine engines typically include a fan and a turbine arranged in flow communication with each other. In addition, turbines of gas turbine engines typically include a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order. In operation, air is provided from the fan to the inlet of the compressor section, wherein one or more axial flow compressors progressively compress the air until it reaches the combustion section. The fuel is mixed with compressed air and combusted within the combustion section to provide combustion gases. The combustion gases are channeled from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed through the exhaust section, for example to the atmosphere. In addition, optimizing blade tip clearances may improve engine performance and efficiency.

Drawings

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary ductless gas turbine engine in accordance with various embodiments of the subject disclosure.

FIG. 2 is a schematic diagram of an exemplary Active Clearance Control (ACC) assembly, according to various embodiments of the present disclosure.

FIG. 3 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 4 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 5 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 6 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 7 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 8 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 9 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 10 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

FIG. 11 is a schematic diagram of another exemplary ACC assembly, according to various embodiments of the present disclosure.

Repeated use of reference characters in the specification and drawings is intended to represent the same or analogous features or elements of the subject matter.

Detailed Description

Reference will now be made in detail to the present embodiments of the disclosure, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar reference numerals have been used in the drawings and the description to refer to the same or similar parts of the disclosure.

The word "exemplary" is used herein to mean "serving as an example, instance, or illustration. Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, all embodiments described herein are to be considered exemplary unless explicitly stated otherwise.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, and are not intended to represent the location or importance of the respective components.

The terms "forward" and "aft" refer to relative positions within a turbine, gas turbine engine, or carrier, and refer to their normal operational attitude. For example, for a gas turbine engine, the front refers to a location closer to the engine inlet and the rear refers to a location closer to the engine nozzle or exhaust.

The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction of fluid flow and "downstream" refers to the direction of fluid flow.

The terms "coupled," "fixed," "attached to" and the like refer to direct coupling, fixing or attaching, as well as indirect coupling, fixing or attaching via one or more intermediate components or features, unless otherwise specified herein.

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

In the context of, for example, "at least one of A, B and C," the term "at least one" refers to a alone, B alone, C alone, or any combination of A, B and C.

Throughout this specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

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

The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid electric versions of one or more of these engines.

In certain aspects of the present disclosure, a three-stream engine includes a turbine for compressing an air stream in a compressor and combusting the compressed air stream to produce a combusted gas. The combusted gases expand in the turbine section. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly surrounding the core engine and at least partially defining a core duct. The fan shroud is disposed radially outward from the core shroud and annularly surrounds at least a portion of the core shroud. The fan housing at least partially defines an inlet duct and a fan duct. The fan duct and the core duct are at least partially co-axially extending on opposite sides of the core cowl. Embodiments of the Active Clearance Control (ACC) assembly of the present disclosure provide cooling air from a fan duct to one or more ACC mechanisms associated with a turbine section of an engine. For example, the clearance between rotating and stationary turbomachine components of the engine may be adjusted by an ACC mechanism. Hot control air may be delivered to the ACC mechanism such that the radial position of the housing and shroud may be adjusted relative to the tips of the rotating blades. According to an exemplary embodiment of the present disclosure, the cooling air provided from the fan duct to the ACC mechanism is typically at a higher pressure. Furthermore, providing cooling air from the fan duct to the ACC mechanism may prevent the cooling air from flowing out of the core airflow, which may otherwise create detrimental combustion losses.

Referring now to FIG. 1, a schematic cross-sectional view of a gas turbine engine 100 is provided according to an example embodiment of the present disclosure. In particular, FIG. 1 provides a turbofan engine having a rotor assembly with single stage ductless rotor blades. In this manner, the rotor assembly may be referred to herein as a "ductless fan," or the entire engine 100 may be referred to as a "ductless turbofan engine. In addition, engine 100 of FIG. 1 includes a third flow extending from the compressor section to a rotor assembly flow path on the turbine, as will be explained in more detail below.

For reference, engine 100 defines an axial direction a, a radial direction R, and a circumferential direction C. Further, engine 100 defines an axial centerline or longitudinal axis 112 extending along axial direction a. In general, the axial direction a extends parallel to the longitudinal axis 112, the radial direction R extends outwardly from the longitudinal axis 112 and inwardly toward the longitudinal axis 112 in a direction orthogonal to the axial direction a, and the circumferential direction extends three hundred sixty degrees (360 °) around the longitudinal axis 112. Engine 100 extends between a forward end 114 and an aft end 116, for example, along an axial direction a.

Engine 100 includes a turbine 120 and a rotor assembly, also referred to as a fan section 150, located upstream thereof. Generally, the turbine 120 includes a compressor section, a combustion section, a turbine section, and an exhaust section in series flow order. Specifically, as shown in FIG. 1, turbine 120 includes a core shroud 122 defining an annular core inlet 124. The core cowl 122 also at least partially encloses the low pressure system and the high pressure system. For example, the illustrated core cowl 122 at least partially encloses and supports a booster or low pressure ("LP") compressor 126 for pressurizing air entering the turbine 120 through the core inlet 124. A high pressure ("HP") multistage axial compressor 128 receives pressurized air from the LP compressor 126 and further increases the pressure of the air. The pressurized air stream flows downstream to the combustor 130 of the combustion section, where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air.

It should be understood that as used herein, the terms "high/low speed" and "high/low pressure" are used interchangeably with respect to high pressure/high speed systems and low pressure/low speed systems. Furthermore, it should be understood that the terms "high" and "low" are used in the same context to distinguish between two systems and are not meant to imply any absolute velocity and/or pressure values.

The high energy combustion products flow downstream from the combustor 130 to a high pressure turbine 132. The high pressure turbine 132 drives the high pressure compressor 128 through a high pressure shaft 136. In this regard, the high pressure turbine 132 is drivingly coupled with the high pressure compressor 128. The high energy combustion products then flow to low pressure turbine 134. Low pressure turbine 134 drives low pressure compressor 126 and components of air sector section 150 via low pressure shaft 138. In this regard, low pressure turbine 134 is drivingly coupled with low pressure compressor 126 and components of air sector section 150. In the exemplary embodiment, LP shaft 138 is coaxial with HP shaft 136. After driving each of the turbines 132, 134, the combustion products exit the turbine 120 through a turbine exhaust nozzle 140.

Thus, the turbine 120 defines a working gas flow path or core duct 142 extending between the core inlet 124 and the turbine exhaust nozzle 140. The core tube 142 is an annular tube positioned generally inside the core housing 122 in the radial direction R. The core conduit 142 (e.g., the working gas flow path through the turbine 120) may be referred to as a second flow.

The fan section 150 includes a fan 152, which in this example embodiment is the main fan. For the embodiment shown in fig. 1, the fan 152 is an open rotor or ductless fan 152. In this manner, engine 100 may be referred to as an open rotor engine.

As shown, the fan 152 includes an array of airfoils disposed about the longitudinal axis 112 of the engine 100, and more specifically, an array of fan blades 154 (only one shown in fig. 1) disposed about the longitudinal axis 112 of the engine 100. The fan blades 154 may, for example, rotate about the longitudinal axis 112. As described above, fan 152 is drivingly coupled with low-pressure turbine 134 via LP shaft 138. For the embodiment shown in FIG. 1, the fan 152 is coupled with the LP shaft 138 via a reduction gearbox 155, such as in an indirect drive or gear drive configuration.

Further, the array of fan blades 154 may be equally spaced about the longitudinal axis 112. Each fan blade 154 has a root and a tip and a span defined therebetween. Each fan blade 154 defines a pitch change or center blade axis 156. For this embodiment, each fan blade 154 of the fan 152 is rotatable about its central blade axis 156, e.g., in unison with each other. One or more actuators 158 are provided to facilitate such rotation, and thus may be used to change the pitch of the fan blades 154 about their respective central blade axes 156.

The fan section 150 also includes an array of airfoils positioned aft of the fan blades 154 and also disposed about the longitudinal axis 112, and more specifically, an array of fan guide vanes 160 that includes fan guide vanes 162 (only one shown in FIG. 1) disposed about the longitudinal axis 112. For this embodiment, the fan guide vanes 162 are not rotatable about the longitudinal axis 112. Each fan guide vane 162 has a root and a tip and a span defined therebetween. The fan guide vanes 162 may be uncovered, as shown in fig. 1, or alternatively, may be covered by, for example, an annular shroud spaced outwardly from the tips of the fan guide vanes 162 along the radial direction R or attached to the fan guide vanes 162.

Each fan guide vane 162 defines a central blade axis 164. For this embodiment, each fan guide vane 162 of the fan guide vane array 160 is rotatable about its respective central vane axis 164, e.g., in unison with each other. One or more actuators 166 are provided to facilitate such rotation and thus may be used to vary the pitch of the fan guide vanes 162 about their respective central blade axes 164. However, in other embodiments, each fan guide vane 162 may be fixed or unable to pitch about its fan guide vane 162. The fan guide vanes 162 are mounted to a fan case 170.

As shown in fig. 1, in addition to ductless fan 152, ducted fan 184 is included aft of fan 152 such that engine 100 includes both ducted and ductless fans for generating thrust by movement of air without a passage through at least a portion of turbine 120 (e.g., without passing through HP compressor 128 and combustion section for the illustrated embodiment). Ducted fan 184 is rotatable about the same axis (e.g., longitudinal axis 112) as fan blades 154. For the depicted embodiment, ducted fan 184 is driven by low pressure turbine 134 (e.g., coupled to LP shaft 138). In the illustrated embodiment, as described above, the fan 152 may be referred to as a primary fan and the duct fan 184 may be referred to as a secondary fan. It should be understood that these terms "primary" and "secondary" are for convenience and are not meant to imply any particular importance, power, etc.

Ducted fan 184 includes a plurality of fan blades (not separately labeled in fig. 1) arranged in a single stage such that ducted fan 184 may be referred to as a single stage fan. The fan blades of ducted fan 184 may be equally spaced about longitudinal axis 112. Each blade of ducted fan 184 has a root and a tip and a span defined therebetween.

The fan shroud 170 annularly surrounds at least a portion of the core shroud 122 and is positioned generally along the radial direction R outside of at least a portion of the core shroud 122. Specifically, a downstream section of the fan shroud 170 extends over a forward portion of the core shroud 122 to define a fan duct flow path or simply define a fan duct 172. According to this embodiment, the fan flow path or fan duct 172 may be understood to form at least a portion of the third flow of the engine 100.

The incoming air may enter the fan duct 172 through a fan duct inlet 176 and may exit through a fan exhaust nozzle 178 to generate propulsive thrust. The fan duct 172 is an annular duct positioned generally outside the core duct 142 along the radial direction R. The fan shroud 170 and the core shroud 122 are coupled together and supported by a plurality of substantially radially extending, circumferentially spaced apart stationary struts 174 (only one shown in FIG. 1). Each stationary strut 174 may have an aerodynamic profile to direct air flow therethrough. Other struts besides the fixed struts 174 may be used to connect and support the fan shroud 170 and/or the core shroud 122. In many embodiments, the fan duct 172 and the core duct 142 may be at least partially coextensive (typically axially) on opposite sides (e.g., opposite radial sides) of the core cowl 122. For example, the fan duct 172 and the core duct 142 may each extend directly from the leading edge 144 of the core cowl 122, and may be partially substantially axially coextensive on opposite radial sides of the core cowl 122.

Engine 100 also defines or includes an inlet duct 180. An inlet duct 180 extends between an engine inlet 182 and the core inlet 124/fan duct inlet 176. The engine inlet 182 is generally defined at the forward end of the fan 170 and is positioned between the fan 152 and the fan guide vane array 160 along the axial direction a. The inlet duct 180 is an annular duct that is positioned inside the fan housing 170 along the radial direction R. The air flowing downstream along the inlet duct is divided (not necessarily uniformly) into core duct 142 and fan duct 172 by fan duct splitter or leading edge 144 of core shroud 122. In the illustrated embodiment, the inlet duct 180 is wider in the radial direction R than the core duct 142. The inlet duct 180 is also wider than the fan duct 172 in the radial direction R.

Notably, for the illustrated embodiment, engine 100 includes one or more features to increase the efficiency of third flow thrust Fn _3S (e.g., thrust generated by the airflow passing through fan duct 172 and exiting through fan exhaust nozzle 178, at least in part, by ducted fan 184). Specifically, engine 100 also includes an array of inlet guide vanes 186 positioned in inlet duct 180 upstream of ducted fan 184 and downstream of engine inlet 182. An array of inlet guide vanes 186 is arranged about the longitudinal axis 112. For this embodiment, the inlet guide vanes 186 cannot rotate about the longitudinal axis 112. Each inlet guide vane 186 defines a central vane axis (not labeled for clarity) and is rotatable about its respective central vane axis, e.g., in unison with each other. In this way, the inlet guide vanes 186 may be considered as variable geometry components. One or more actuators 188 are provided to facilitate such rotation, and thus may be used to vary the pitch of the inlet guide vanes 186 about their respective central blade axes. However, in other embodiments, each inlet guide vane 186 may be fixed or unable to pitch about its central vane axis.

Further, downstream of ducted fan 184 and upstream of duct inlet 176, engine 100 includes an array of outlet guide vanes 190. As with the array of inlet guide vanes 186, the array of outlet guide vanes 190 cannot rotate about the longitudinal axis 112. However, for the depicted embodiment, unlike the array of inlet guide vanes 186, the array of outlet guide vanes 190 is configured as fixed pitch outlet guide vanes.

Furthermore, it should be appreciated that for the depicted embodiment, the fan exhaust nozzle 178 of the fan duct 172 is also configured as a variable geometry exhaust nozzle. In this manner, engine 100 includes one or more actuators 192 for adjusting the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to vary the total cross-sectional area (e.g., the area of the nozzle in a plane perpendicular to the longitudinal axis 112) to adjust the amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flow, etc., of the airflow through the fan duct 172). A fixed geometry exhaust nozzle may also be employed.

Further, still referring to FIG. 1, in the exemplary embodiment, air passing through fan duct 172 may be relatively cooler (e.g., lower temperature) than one or more fluids used in turbine 120. In this manner, one or more heat exchangers 200 may be positioned in thermal communication with the fan duct 172. For example, one or more heat exchangers 200 may be disposed within the fan duct 172 and configured to cool one or more fluids from the core engine with air flowing through the fan duct 172 as a resource for removing heat from the fluids (e.g., compressor bleed air, oil, or fuel).

Referring now to fig. 2, fig. 2 is a schematic illustration of an embodiment of an Active Clearance Control (ACC) assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the ACC assembly 210 includes an active clearance control mechanism 212 operatively associated with a turbine section 214 of the gas turbine engine 100. For example, as shown in FIG. 1, the turbine section 214 includes a high pressure turbine 132 and a low pressure turbine 134. The high pressure turbine 132 and the low pressure turbine 134 may include one or more shroud assemblies (not shown), each shroud assembly forming an annular ring surrounding an annular array of rotor blades of the respective high pressure turbine 132 and low pressure turbine 134. The shroud assembly may be coupled with a hanger (not shown) that is in turn coupled with a corresponding high pressure turbine housing 220 and a corresponding low pressure turbine housing 222. Generally, the shroud of the shroud assembly is radially spaced from the blade tips of each of the array of rotor blades 226 of the high pressure turbine 132 and the array of rotor blades 228 of the low pressure turbine 134. The shroud generally reduces clearances and leakage across the blade tips in order to maximize turbine power extracted from the core airflow through the core duct 142 via the rotor blades 226 and 228. A blade tip clearance is generally defined between the blade tip and the shroud. It should be appreciated that engine performance parameters (e.g., thrust, specific Fuel Consumption (SFC), exhaust Gas Temperature (EGT), emissions, etc.) depend, at least in part, on the clearance between the turbine blade tips and the shroud of the shroud assembly. The clearance between the turbine blade tips and the shroud is typically minimized to promote optimal engine performance and efficiency. The challenge in minimizing the clearances is that mechanical and thermal loads acting on the turbomachine component during engine operation cause the component to expand and contract at different rates. For example, the rotor and the casing surrounding the blades contract and expand at different rates.

Thus, the ACC assembly 210 is a system that controls and optimizes clearance throughout various phases of the flight. As will be appreciated, ACC assembly 210 conditions a relatively cool or hot air flow from the source of gas turbine engine 100 and distributes the air over the HP and/or LP turbine housings and shrouds to retract or expand the engine housing relative to the turbine blade tips, depending on factors such as the operation and flight conditions of the aircraft. In this way, the clearance may be adjusted to optimize engine performance.

In the illustrated embodiment, ACC mechanism 212 includes an HP turbine ACC mechanism 240 and an LP turbine ACC mechanism 242. The HP turbine ACC mechanism 240 controls and optimizes the clearances associated with the HP turbine 132, and the LP turbine ACC mechanism 242 controls and optimizes the clearances associated with the LP turbine 134. The ACC flow path 252 is defined by the ACC assembly 210 and is a flow path for fluid flow (e.g., bleed air or extraction air) from the fan duct 172 to and/or through components of the ACC assembly 210.

In FIG. 2, ACC assembly 210 includes a duct assembly 250, duct assembly 250 forming part of an ACC flow path 252 and thermally coupled to fan duct 172 and ACC mechanism 212. In fig. 2, the duct assembly 250 includes an air supply inlet 254 fluidly connected to the fan duct 172 and downstream of the heat exchanger 200 to draw a portion of a fan duct airflow 256 through the fan duct 172. The conduit assembly 250 includes a flow control device 258 thermally coupled to the ACC flow path 252 and the air supply inlet 254. The flow control device 258 is fluidly connected to the air supply inlet 254 via a line 260. The line 260 partially defines the ACC flow path 252. A line 260 is also fluidly connected to the air supply inlet 254 and extends from the air supply inlet 254 such that, in the ACC flow path 252, the flow control device 258 is located downstream of the air supply inlet 254. In the illustrated embodiment, line 262 is fluidly connected to LP turbine ACC mechanism 242. Flow control device 258 regulates and/or controls the flow of air through conduit assembly 250 to ACC mechanism 212. For example, in the illustrated embodiment, the flow control device 258 may be a valve that regulates the flow of air delivered to the LP turbine ACC mechanism 242.

During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or conveyed as an ACC fluid stream 264 into the air supply inlet 254. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. The LP turbine ACC mechanism 242 uses the ACC fluid flow 264 to control and optimize the clearances associated with the LP turbine 134 (e.g., a system utilizing manifolds, plenums, etc.).

Referring now to fig. 3, fig. 3 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, ACC assembly 210 is configured similarly to the embodiment shown in FIG. 2, except that line 262 is fluidly connected to HP turbine ACC mechanism 240. During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or routed into the air supply inlet 254 as an ACC fluid flow 264. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the HP turbine ACC mechanism 240. The HP turbine ACC mechanism 240 uses the ACC fluid flow 264 to control and optimize clearances associated with the HP turbine 132 (e.g., systems utilizing manifolds, plenums, etc.).

Embodiments of the present disclosure use cooling air for the LP turbine ACC mechanism 242 and/or the HP turbine ACC mechanism 240 that has been used as a radiator in the heat exchanger 200. Furthermore, because cooling air for the LP turbine ACC mechanism 242 and/or the HP turbine ACC mechanism 240 is drawn from the fan duct 172 instead of the core duct 142 (or fan flow), at lower power and/or cruise conditions, thrust is not affected during peak power conditions.

Referring now to fig. 4, fig. 4 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, ACC assembly 210 is configured similarly to the embodiment shown in fig. 2, except that air supply inlet 254 is fluidly connected to fan duct 172 at a location upstream of heat exchanger 200. During operation of gas turbine engine 100, a portion of fan duct airflow 256 is extracted from fan duct 172 upstream of heat exchanger 200 (e.g., by a scoop or other type of mechanism) and channeled or routed as ACC fluid flow 264 into air supply inlet 254. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. The LP turbine ACC mechanism 242 uses the ACC fluid flow 264 to control and optimize clearances associated with the LP turbine 134.

Referring now to fig. 5, fig. 5 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, ACC assembly 210 is similarly configured to the embodiment shown in FIG. 4, except that line 262 is fluidly connected to HP turbine ACC mechanism 240. During operation of gas turbine engine 100, a portion of fan duct airflow 256 is extracted from fan duct 172 upstream of heat exchanger 200 (e.g., by a scoop or other type of mechanism) and channeled or routed as ACC fluid flow 264 into air supply inlet 254. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the HP turbine ACC mechanism 240. The HP turbine ACC mechanism 240 uses the ACC fluid flow 264 to control and optimize clearances associated with the HP turbine 132.

Referring now to FIG. 6, FIG. 6 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the conduit assembly 250 of the ACC assembly 210 includes a conduit assembly 250A forming an ACC flow path 252A and a conduit assembly 250B forming an ACC flow path 252B. ACC flow paths 252A and 252B are thermally connected to fan duct 172.

In the illustrated embodiment, the air supply inlet 270 is fluidly connected to the fan duct 172 and is located downstream of the heat exchanger 200 to draw a portion of the fan duct airflow 256 through the fan duct 172. The conduit assembly 250A includes a flow control device 272 that is thermally coupled to the ACC flow path 252A and the air supply inlet 270. The flow control device 272 is fluidly connected to the air supply inlet 270 via line 274. Line 274 partially defines ACC flow path 252A. The line 274 is also fluidly connected to the air supply inlet 270 and extends from the air supply inlet 270 such that, in the ACC flow path 252A, the flow control device 272 is located downstream of the air supply inlet 270. The flow control device 272 is fluidly connected to the HP turbine ACC mechanism 240 via line 276. Line 276 partially defines ACC flow path 252A. Line 276 is also fluidly connected to HP turbine ACC mechanism 240 and extends to HP turbine ACC mechanism 240 such that, in ACC flow path 252A, HP turbine ACC mechanism 240 is downstream of flow control device 272. The flow control device 272 regulates and/or controls the ACC fluid flow 264A through the conduit assembly 250A to the HP turbine ACC mechanism 240. For example, in the illustrated embodiment, the flow control device 272 may be a valve that regulates the flow of air to the HP turbine ACC mechanism 240.

In fig. 6, duct assembly 250B includes an air supply inlet 280 fluidly connected to fan duct 172 and downstream of heat exchanger 200 to draw a portion of fan duct airflow 256 through fan duct 172. The conduit assembly 250B includes a flow control device 282 that is thermally coupled to the ACC flow path 252B and the air supply inlet 280. The flow control device 282 is fluidly connected to the air supply inlet 280 via a line 284. Line 284 partially defines ACC flow path 252B. The line 284 is also fluidly connected to the air supply inlet 280 and extends from the air supply inlet 280 such that, in the ACC flow path 252B, the flow control device 282 is located downstream of the air supply inlet 280. The flow control device 282 is fluidly connected to the LP turbine ACC mechanism 242 via a line 286. Line 286 partially defines ACC flow path 252B. The line 286 is also fluidly connected to the LP turbine ACC mechanism 242 and extends to the LP turbine ACC mechanism 242 such that, in the ACC flow path 252B, the LP turbine ACC mechanism 242 is downstream of the flow control apparatus 282. The flow control device 282 regulates and/or controls the ACC fluid flow 264B through the conduit assembly 250B to the LP turbine ACC mechanism 242. For example, in the illustrated embodiment, the flow control device 282 may be a valve that regulates the flow of air delivered to the LP turbine ACC mechanism 242. In the illustrated embodiment, an air supply inlet 280 corresponding to the LP turbine ACC mechanism 242 is located within the fan duct 172 downstream of the air supply inlet 270 associated with the HP turbine ACC mechanism 240. However, it should be appreciated that the upstream/downstream positions of the air supply inlets 270 and 280 may be reversed.

During operation of gas turbine engine 100, a portion of fan duct airflow 256 is extracted from fan duct 172 downstream of heat exchanger 200 (e.g., by a scoop or other type of mechanism) and channeled or conveyed as ACC fluid streams 264A and 264B, respectively, into air supply inlets 270 and/or 280. The flow control devices 272 and 282 control the volumes of ACC fluid flows 264A and 264B delivered to the HP turbine ACC mechanism 240 and the LP turbine ACC mechanism 242, respectively. The HP turbine ACC mechanism 240 uses the ACC fluid flow 264A to control and optimize the clearance associated with the HP turbine 132, and the LP turbine ACC mechanism 242 uses the ACC fluid flow 264B to control and optimize the clearance associated with the LP turbine 134.

Referring now to FIG. 7, FIG. 7 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, ACC assembly 210 is configured similarly to the embodiment shown in fig. 6, except that air supply inlet 280 is omitted and line 284 of conduit assembly 250B is fluidly connected to line 274. During operation of gas turbine engine 100, a portion of fan duct airflow 256 is extracted from fan duct 172 downstream of heat exchanger 200 (e.g., by a scoop or other type of mechanism) and directed or routed into air supply inlet 270. The flow control devices 272 and 282 control the volume or flow rate of the extracted airflow to deliver the extracted airflow as ACC fluid flows 264A and 264B, respectively. The flow control devices 272 and 282 control the volumes of ACC fluid flows 264A and 264B delivered to the HP turbine ACC mechanism 240 and the LP turbine ACC mechanism 242, respectively. The HP turbine ACC mechanism 240 uses the ACC fluid flow 264A to control and optimize the clearance associated with the HP turbine 132, and the LP turbine ACC mechanism 242 uses the ACC fluid flow 264B to control and optimize the clearance associated with the LP turbine 134.

Referring now to fig. 8, fig. 8 is a schematic view of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the ACC assembly 210 is configured similarly to the embodiment depicted in fig. 2, with the air supply inlet 254 being located downstream of the heat exchanger 200 in the fan duct 172, and the air supply inlet 254 being fluidly connected to a flow control device 258 in the ACC flow path 252 downstream of the air supply inlet 254. In FIG. 8, the flow control device 258 is fluidly connected to the LP turbine ACC mechanism 242 via a line 290. The line 290 partially defines a first portion 292 of the ACC fluid flow 252 such that the line 290 is fluidly connected to the LP turbine ACC mechanism 242. In the first portion 292 of the ACC flow path 252, the LP turbine ACC mechanism 242 is located downstream of the flow control apparatus 258. Downstream of the flow control device 258 in the ACC flow path 252, a line 294 is fluidly connected to a line 290. The line 294 partially defines a second portion 296 of the ACC flow path 252 such that the line 294 is fluidly connected to the HP turbine ACC mechanism 240, wherein in the second portion 296 of the ACC flow path 252 the HP turbine ACC mechanism 240 is downstream of the flow control device 258.

During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or routed as an ACC fluid stream 264 to the air supply inlet 254. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the HP turbine ACC mechanism 240 and the LP turbine ACC mechanism 242.

Referring now to fig. 9, fig. 9 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the ACC assembly 210 is configured similarly to the embodiment depicted in fig. 2, with the air supply inlet 254 being located downstream of the heat exchanger 200 in the fan duct 172, and the air supply inlet 254 being fluidly connected to a flow control device 258 in the ACC flow path 252 downstream of the air supply inlet 254.

In the illustrated embodiment, the flow control device 258 is fluidly connected to the core duct 142 at a compressor section of the turbine 120. For example, in the illustrated embodiment, the flow control device 258 is fluidly connected to the core duct 142 at an air supply inlet 300, the air supply inlet 300 being located axially downstream of the LP compressor 126 and upstream of the HP compressor 128. In some embodiments, the air supply inlet 300 is located after the last stage of the LP compressor 126. However, it should be appreciated that the air supply inlet 300 may be located elsewhere within the compressor section of the turbine 120. The flow control device 258 is fluidly connected to the air supply inlet 300 by a line 302. The line 302 partially defines the ACC flow path 252 such that the line 302 is fluidly connected to the flow control device 258 such that in the ACC flow path 252, the flow control device 258 is located downstream of the air supply inlet 300.

The flow control device 258 controls and/or regulates the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. In addition, the flow control device 258 controls and/or regulates the source of the ACC fluid flow 264. For example, the flow control device 258 may be a valve operable to select a source of ACC fluid flow as the fan duct 172 or the core duct 142. During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or routed into the air supply inlet 254 as an ACC fluid flow 264. Alternatively, a portion of the core airflow 230 is extracted from the core duct 142 (e.g., by a scoop or other type of mechanism) and directed or conveyed as an ACC fluid flow 264 into the air supply inlet 300. The flow control device 258 controls the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. The LP turbine ACC mechanism 242 uses the ACC fluid flow 264 to control and optimize clearances associated with the LP turbine 134.

Referring now to fig. 10, fig. 10 is a schematic view of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the ACC assembly 210 is configured similarly to the embodiment depicted in fig. 2, with the air supply inlet 254 downstream of the heat exchanger 200 in the fan duct 172, and in the ACC flow path 252, the air supply inlet 254 is fluidly connected to the flow control device 258 downstream of the air supply inlet 254 via a line 260.

In the illustrated embodiment, the flow control device 304 is fluidly connected to the core tube 142 at a compressor section of the turbine 120. For example, in the illustrated embodiment, flow control device 304 is fluidly connected to core tube 142 at air supply inlet 300, with air supply inlet 300 being located axially downstream of LP compressor 126 and upstream of HP compressor 128. In some embodiments, the air supply inlet 300 is located after the last stage of the LP compressor 126. However, it should be appreciated that the air supply inlet 300 may be located at other locations within the compressor section of the turbine 120. The flow control device 304 is fluidly connected to the air supply inlet 300 by a line 308. The line 308 partially defines the ACC flow path 252 such that the line 308 is fluidly connected to the flow control device 304 such that, in the ACC flow path 252, the flow control device 304 is downstream of the air supply inlet 300. Flow control device 304 is fluidly connected to line 262 by line 310. Line 310 partially defines ACC flow path 252 such that line 310 is fluidly connected to flow control device 304 such that, in ACC flow path 252, line 310 is downstream of flow control device 304. Further, in ACC flow path 252, downstream of flow control devices 258 and 304, line 310 is fluidly connected to line 262.

The flow control devices 258 and 304 control and/or regulate the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. In addition, the flow control devices 258 and 304 control and/or regulate the source of the ACC fluid flow 264. For example, the flow control device 258 may be a valve operable to select the source of the ACC fluid flow 264 as the fan duct 172 and the flow control device 304 may be a valve operable to select the source of the ACC fluid flow 264 as the core duct 142. During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or routed into the air supply inlet 254 as an ACC fluid flow 264. Alternatively or in combination therewith, a portion of the core airflow 230 is extracted from the core duct 142 (e.g., by a scoop or other type of mechanism) and directed or conveyed as an ACC fluid stream 264 into the air supply inlet 300. The flow control device 304 controls the volume 242 of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism. The LP turbine ACC mechanism 242 uses the ACC fluid flow 264 to control and optimize clearances associated with the LP turbine 134.

Referring now to FIG. 11, FIG. 11 is a schematic illustration of another embodiment of an ACC assembly 210 for a gas turbine engine 100 according to the present disclosure. The gas turbine engine 100 may be configured in a similar manner as the exemplary gas turbine engine 100 of FIG. 1. In the illustrated embodiment, the ACC assembly 210 is configured similarly to the embodiment depicted in fig. 10, with the air supply inlet 254 downstream of the heat exchanger 200 in the fan duct 172, and in the ACC flow path 252, the air supply inlet 254 is fluidly connected to the flow control device 258 downstream of the air supply inlet 254 via a line 260. Also, the flow control apparatus 304 is fluidly connected to the core tube 142 at the compressor section of the turbine 120. The flow control device 304 is fluidly connected to the core duct 142 at an air supply inlet 300, the air supply inlet 300 being located axially downstream of the LP compressor 126 and upstream of the HP compressor 128. In the illustrated embodiment, the air supply inlet 300 is located after the last stage of the LP compressor 126. However, it should be understood that the air supply inlet 300 may be located elsewhere in the compressor section of the turbine 120. The flow control device 304 is fluidly connected to the air supply inlet 300 by a line 308 such that in the ACC flow path 252, the flow control device 304 is downstream of the air supply inlet 300. Flow control device 304 is fluidly connected to line 262 by line 310. The line 310 is fluidly connected to the flow control device 304 such that in the ACC flow path 252, the line 310 is downstream of the flow control device 304. In the illustrated embodiment, in the ACC flow path 252, upstream of the flow control device 258, a line 310 is fluidly connected to the line 260.

The flow control devices 258 and 304 control and/or regulate the volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242. In addition, the flow control devices 258 and 304 control and/or regulate the source of ACC. For example, the flow control device 304 may be a valve operable to select a source of ACC fluid flow 264 as the core conduit 142. The fluid flow from core conduit 142 may be mixed with the fluid flow from fan conduit 172 before reaching flow control device 258. The fluid control device 258 controls and/or regulates a volume of the ACC fluid flow 264 delivered to the LP turbine ACC mechanism 242.

During operation of the gas turbine engine 100, a portion of the fan duct airflow 256 is extracted from the fan duct 172 (e.g., by a scoop or other type of mechanism) and directed or routed into the air supply inlet 254 as an ACC fluid flow 264. A portion of the core airflow 230 may be extracted from the core duct 142 (e.g., by a scoop or other type of mechanism) and directed or conveyed as an ACC fluid flow 264 into the air supply inlet 300. Actuation of the flow control device 304 directs a portion of the core airflow 230 received by the air supply inlet 300 to flow downstream and mix with the fluid flow received from the air supply inlet 254. The flow control device 258 then controls the volume of the ACC fluid flow 264 generated by the mixed fluid flow delivered to the LP turbine ACC mechanism 242. The LP turbine ACC mechanism 242 uses the ACC fluid flow 264 to control and optimize clearances associated with the LP turbine 134.

In the embodiment depicted and described in fig. 9-11, ACC fluid stream 264 is delivered to LP turbine ACC mechanism 242. It should be appreciated that in the embodiment illustrated in FIGS. 9-11, ACC fluid flow 264 may alternatively or additionally be provided to HP turbine ACC mechanism 240.

Accordingly, embodiments of the Active Clearance Control (ACC) assembly of the present disclosure provide cooling air from the fan duct to one or more ACC mechanisms associated with the turbine section of the engine. For example, according to an exemplary embodiment of the present disclosure, cooling air provided from the fan duct to the ACC mechanism is typically at a higher pressure. Furthermore, providing cooling air from the fan duct to the ACC mechanism may prevent the cooling air from flowing out of the core airflow, which may otherwise create detrimental combustion losses.

This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

While this disclosure has been described as having an exemplary design, the present disclosure may be further modified within the scope of this disclosure. This application is therefore intended to cover any variations, uses, or adaptations of the disclosure using its general principles. Furthermore, this application is intended to cover such departures from the present disclosure as come within known or customary practice in the art to which this disclosure pertains and which fall within the limits of the appended claims.

Further aspects are provided by the subject matter of the following clauses:

a gas turbine engine, comprising: a fan section, an engine inlet, and a fan duct splitter in serial flow order, the fan duct splitter splitting an airflow entering the engine inlet from the fan section to a fan duct and a core duct, the core duct comprising a compressor section, a combustion section, and a turbine section in serial flow; and a duct assembly coupled to the fan duct to extract a portion of the fan duct airflow through the fan duct and to deliver the portion of the fan duct airflow to the active clearance control mechanism of the turbine section.

The gas turbine engine of any preceding claim, wherein the fan duct comprises a heat exchanger, and wherein the duct assembly is coupled to the fan duct downstream of the heat exchanger.

The gas turbine engine of any preceding claim, wherein the fan duct comprises a heat exchanger, and wherein the duct assembly is coupled to the fan duct upstream of the heat exchanger.

The gas turbine engine of any of the preceding clauses, further comprising a flow control device that regulates a flow of the portion of the fan duct airflow to the active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, wherein the turbine section comprises a low pressure turbine, and wherein the active clearance control mechanism is operably coupled with the low pressure turbine.

The gas turbine engine of any of the preceding clauses, wherein the turbine section comprises a high pressure turbine, and wherein the active clearance control mechanism is operably coupled with the high pressure turbine.

The gas turbine engine of any of the preceding clauses, wherein the turbine section comprises a low pressure turbine and a high pressure turbine, and wherein the active clearance control mechanism comprises a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, further comprising a flow control device that regulates a flow of a portion of the fan duct airflow to the low pressure turbine active clearance control mechanism and the high pressure turbine active clearance control mechanism.

The gas turbine engine of any of the preceding claims, wherein the duct assembly is further coupled to the core duct to extract a portion of the core airflow through the core duct and to deliver a portion of the core airflow to the active clearance control mechanism of the turbine section.

The gas turbine engine of any of the preceding clauses, further comprising a flow control device that regulates a flow of at least one of a portion of the fan duct airflow or a portion of the core airflow delivered to the active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, wherein the active clearance control mechanism comprises a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism, and wherein the duct assembly defines a first active clearance control flow path from the fan duct to the low pressure turbine active clearance control mechanism and a second active clearance control flow path from the fan duct to the high pressure turbine active clearance control mechanism.

The gas turbine engine according to any one of the preceding clauses, further comprising a first flow control device within the first active clearance control flow path and a second flow control device within the second active clearance control flow path.

The gas turbine engine of any of the preceding claims, wherein the active clearance control mechanism comprises a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism, and wherein the duct assembly comprises an air supply inlet in communication with the fan duct to supply a portion of the fan duct airflow along a first active clearance control flow path from the air supply inlet to the low pressure turbine active clearance control mechanism and a second active clearance control flow path from the air supply inlet to the high pressure turbine active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, further comprising a first flow control device within a first active clearance control flow path between the air supply inlet and the low pressure turbine active clearance control mechanism and a second flow control device within a second active clearance control flow path between the air supply inlet and the high pressure turbine active clearance control mechanism.

The gas turbine engine of any preceding claim, wherein the duct assembly comprises a first air supply inlet in communication with the fan duct and a second air supply inlet in communication with the core duct.

The gas turbine engine of any preceding claim, wherein the duct assembly defines an active clearance control flow path from the first air supply inlet to the active clearance control mechanism and from the second air supply inlet to the active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, further comprising a first flow control device located between the first air supply inlet and the active clearance control mechanism and a second flow control device located between the second air supply inlet and the active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, wherein the duct assembly comprises a first air supply inlet in communication with the fan duct and a second air supply inlet in communication with the core duct, and wherein the duct assembly comprises a first flow control device downstream of the second air supply inlet and a first flow control device and a second flow control device downstream of the first air supply inlet.

The gas turbine engine of any one of the preceding clauses, wherein the first flow control device regulates a portion of the core airflow delivered to the second flow control device from the second air supply inlet, and wherein the second flow control device regulates a portion of the fan duct airflow received from the first air supply inlet and delivers a portion of the core airflow to the active clearance control mechanism.

A method of providing clearance control for a gas turbine engine having a fan section, an engine inlet, and a fan duct splitter in serial flow order, the fan duct splitter splitting airflow from the fan section into the engine inlet to a fan duct and a core duct, the core duct including a compressor section, a combustion section, and a turbine section in serial flow order, the method comprising: drawing a portion of the fan duct airflow through the fan duct; a portion of the fan duct airflow is transferred to an active clearance control mechanism of the turbine section.

The method of any of the preceding clauses, further comprising adjusting a flow rate of a portion of the fan duct airflow to the active clearance control mechanism.

The method of any of the preceding clauses, wherein the fan duct comprises a heat exchanger, and wherein extracting a portion of the fan duct airflow comprises extracting a portion of the fan duct airflow from a location of the fan duct downstream of the heat exchanger.

The method of any of the preceding clauses, wherein the fan duct comprises a heat exchanger, and wherein extracting a portion of the fan duct airflow comprises extracting a portion of the fan duct airflow from a location of the fan duct upstream of the heat exchanger.

The method of any of the preceding clauses, wherein the turbine section includes a low pressure turbine and a high pressure turbine, and wherein the active clearance control mechanism includes a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism, further comprising adjusting a flow of a portion of the fan duct airflow to the low pressure turbine active clearance control mechanism and the high pressure turbine active clearance control mechanism.

The method of any of the preceding clauses, wherein the turbine section includes a low pressure turbine and a high pressure turbine, and wherein the active clearance control mechanism includes a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism, further comprising transferring a first portion of a portion of the fan duct airflow extracted from the fan duct to the low pressure turbine active clearance control mechanism, and transferring a second portion of the fan duct airflow extracted from the fan duct to the high pressure turbine active clearance control mechanism.

A gas turbine engine, comprising: a core shroud supporting a turbine including a compressor section, a combustion section, and a turbine section arranged in series flow order; a fan assembly rotatable by the turbine; a fan shroud surrounding at least a portion of the core shroud and defining a fan duct extending between the fan shroud and the core shroud; an active clearance control mechanism operatively coupled with the turbine section; and a duct assembly coupled to the fan duct to extract a portion of the fan duct airflow through the fan duct and deliver a portion of the fan duct airflow to the active clearance control mechanism.

The gas turbine engine of any of the preceding clauses, further comprising a heat exchanger disposed within the fan duct, and wherein the duct assembly is operatively coupled to the fan duct upstream of the heat exchanger.

The gas turbine engine of any of the preceding clauses, further comprising a heat exchanger disposed within the fan duct, and wherein the duct assembly is operatively coupled to the fan duct downstream of the heat exchanger.

The gas turbine engine of any of the preceding clauses, further comprising a flow control device that regulates a flow of a portion of the fan duct airflow to the active clearance control mechanism.

Claims (10)

1. A gas turbine engine, comprising:

a fan section, an engine inlet, and a fan duct splitter in serial flow order, the fan duct splitter splitting airflow from the fan section into the engine inlet to a fan duct and a core duct, the core duct comprising a compressor section, a combustion section, and a turbine section in serial flow order; and

A duct assembly coupled to the fan duct to draw a portion of a fan duct airflow through the fan duct and deliver the portion of the fan duct airflow to an active clearance control mechanism of the turbine section.

2. The gas turbine engine of claim 1, wherein the fan duct includes a heat exchanger, and wherein the duct assembly is coupled to the fan duct downstream of the heat exchanger.

3. The gas turbine engine of claim 1, wherein the fan duct includes a heat exchanger, and wherein the duct assembly is coupled to the fan duct upstream of the heat exchanger.

4. The gas turbine engine of claim 1, further comprising a flow control device that regulates a flow of the portion of the fan duct airflow to the active clearance control mechanism.

5. The gas turbine engine of claim 1, wherein the turbine section comprises a low pressure turbine, and wherein the active clearance control mechanism is operatively coupled with the low pressure turbine.

6. The gas turbine engine of claim 1, wherein the turbine section comprises a high pressure turbine, and wherein the active clearance control mechanism is operatively coupled with the high pressure turbine.

7. The gas turbine engine of claim 1, wherein the turbine section includes a low pressure turbine and a high pressure turbine, and wherein the active clearance control mechanism includes a low pressure turbine active clearance control mechanism and a high pressure turbine active clearance control mechanism.

8. The gas turbine engine of claim 7, further comprising a flow control device that regulates flow to the portion of the fan duct airflow to the low pressure turbine active clearance control mechanism and the high pressure turbine active clearance control mechanism.

9. The gas turbine engine of claim 1, wherein the duct assembly is further coupled to the core duct to extract a portion of a core airflow through the core duct and deliver the portion of the core airflow to the active clearance control mechanism of the turbine section.

10. The gas turbine engine of claim 9, further comprising a flow control device that regulates a flow of at least one of the portion of the fan duct airflow or the portion of the core airflow delivered to the active clearance control mechanism.