CN116950772A - Heat transfer system for a gas turbine engine - Google Patents

Abstract

A turbofan engine is provided. The turbofan engine includes: a fan; a turbine operably coupled to the fan to drive the fan, wherein the turbine, the fan, or both comprise an engine component; a heat source; and a heat transfer system configured to reduce ice accumulation or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising: a first heat transfer member in communication with the heat source; and a second heat transfer member extending from the first heat transfer member to or through the engine member, wherein the first heat transfer member comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer member comprises the other of a heat pipe or a graphene rod.

Description

Heat transfer system for a gas turbine engine

PRIORITY INFORMATION

The present application claims priority from indian provisional patent application No. 202211024723 filed on 4/27 of 2022.

Technical Field

The present subject matter relates generally to gas turbine engines, or more particularly, to heat transfer systems configured to reduce ice build-up or ice formation on components of the engine.

Background

Turbofan engines typically include a fan and a turbine having a plurality of fan blades disposed in flow communication with one another. In addition, turbofan 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 where one or more axial 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 directed 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 an exhaust section, for example, to the atmosphere. However, during severe weather, frost rain, hail, rain and snow, ice, etc. may accumulate on the inlet components of the turbofan engine.

Drawings

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

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to an exemplary embodiment of the present subject matter.

FIG. 2 is a close-up schematic cross-sectional view of a portion of the exemplary gas turbine engine of FIG. 1 including an exemplary heat transfer system in accordance with an exemplary embodiment of the present subject matter.

FIG. 3 is a close-up schematic cross-sectional view of a portion of an exemplary gas turbine engine including an exemplary heat transfer system according to another exemplary embodiment of the present subject matter.

FIG. 4 is a close-up schematic cross-sectional view of a portion of an exemplary gas turbine engine including an exemplary heat transfer system according to another exemplary embodiment of the present subject matter.

FIG. 5 provides a block diagram of a control system for controlling a gas turbine engine including a heat transfer system according to an exemplary embodiment of the present disclosure.

FIG. 6 is a flowchart of an exemplary method of providing heat from a gearbox of a turbofan engine to a first engine component and a second engine component of the turbofan engine to reduce ice accretion or ice formation in accordance with an exemplary embodiment of the present disclosure.

FIG. 7 is an example computing system according to an example embodiment of the disclosure.

Corresponding reference characters indicate corresponding parts throughout the several views. The exemplifications set out herein illustrate exemplary embodiments of the disclosure, and such exemplifications are not to be construed as limiting the scope of the disclosure in any manner.

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. Like or similar reference numerals have been used in the drawings and description to refer to like or similar parts of the disclosure.

The following description is presented to enable one of ordinary skill in the art to make and use the described embodiments as contemplated for carrying out the present disclosure. Various modifications, equivalents, changes, and alternatives will be apparent to those skilled in the art. Any and all such modifications, variations, equivalents, and alternatives are intended to fall within the scope of the present disclosure.

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

For purposes of the following description, the terms "vertical," "horizontal," "longitudinal," and derivatives thereof shall relate to the present disclosure as it is oriented in the drawings. However, it should be understood that the present disclosure may assume various alternative variations, except where expressly specified to the contrary. It is also to be understood that the specific devices illustrated in the attached drawings and described in the following specification are simply exemplary embodiments of the disclosure. Accordingly, specific dimensions and other physical characteristics related to the embodiments disclosed herein are not to be considered as limiting.

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 gas turbine engine, where forward refers to a position closer to an engine inlet and aft refers to a position closer to an 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 from which fluid flows and "downstream" refers to the direction in which fluid flows.

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

Furthermore, unless otherwise indicated, the terms "low", "high" or their respective comparison stages (e.g., lower, higher, where applicable) each refer to a relative speed or pressure within the engine. For example, a "low pressure turbine" operates at a pressure that is typically lower than a "high pressure turbine". Alternatively, the above terms may be understood at their highest level unless otherwise indicated. For example, a "low pressure turbine" may refer to the lowest maximum pressure turbine within the turbine section, while a "high pressure turbine" may refer to the highest maximum pressure turbine within the turbine section. The engine of the present disclosure may also include an intermediate pressure turbine, such as an engine having three spools.

Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, values modified by terms such as "about," "approximately," and "substantially" are not limited to the precise values specified. In at least some cases, the approximating language may correspond to the precision of an instrument for measuring the value or the precision of a method or machine for constructing or manufacturing a component and/or system. For example, approximating language may refer to being within a margin of 1%, 2%, 4%, 10%, 15%, or 20%. These approximation margins may be applied to individual values, margins defining either or both endpoints of a numerical range, and/or ranges between endpoints.

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

As described above, during severe weather, frost rain, hail, rain and snow, ice, etc. may accumulate on the inlet components of the turbofan engine. When ice accumulates, it can fall off and be sucked into the engine. In addition, the ice chunks may damage the fan blades or other downstream components of the engine and may cause the engine to stall.

The present disclosure relates generally to a heat transfer system coupled to or integrated into a portion of a turbofan engine. The heat transfer system of the present disclosure is configured to reduce ice accretion or ice formation in engine components of a turbofan engine. The heat transfer system of the present disclosure includes a first heat transfer member in communication with a heat source and a second heat transfer member extending from the first heat transfer member to or through an engine member. The first heat transfer member comprises one of a heat pipe or a graphene rod, and the second heat transfer member comprises the other of a heat pipe or a graphene rod.

In the exemplary embodiment, the heat source is a gearbox of a turbofan engine. In this way, the heat transfer system of the present disclosure utilizes waste heat from the gearbox to reduce ice build-up or ice formation in components of the gas turbine engine.

Different groups of engine components are susceptible to icing under different conditions. Thus, the control system of the present disclosure utilizes the first type of detection information for the first set of engine components and the second type of detection information for the second set of engine components. In this way, the system of the present disclosure allows for proper ice monitoring and detection for each set of engine components. Further, in this manner, the control system of the present disclosure is able to activate the heat transfer system only for the group of engine components that are experiencing a potential icing condition.

Referring now to FIG. 1, a schematic cross-sectional view of a gas turbine engine or turbofan engine (referred to herein as engine 100) is provided in accordance with an example embodiment of the present disclosure. In particular, FIG. 1 provides an engine having a rotor assembly with a single stage non-ducted rotor blade. In this manner, the rotor assembly may be referred to herein as a "non-ducted fan," or the entire engine 100 may be referred to as a "non-ducted engine" or an engine having an open rotor propulsion system 102. Further, engine 100 of FIG. 1 includes a fan flow of a fan assembly flow path extending from the compressor section to a rotor assembly on the turbine, as will be explained in more detail below. It is also contemplated that in other exemplary embodiments, the present disclosure is compatible with engines having ducts surrounding rotor assemblies (also referred to as fan sections). It is also contemplated that in other exemplary embodiments, the present disclosure is compatible with turbofan engines having a third flow as described herein.

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. Typically, the axial direction a extends parallel to the longitudinal axis 112, the radial direction R extends outwardly from the longitudinal axis 112 and inwardly to the longitudinal axis 112 in a direction perpendicular 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, in an axial direction a.

Engine 100 includes a turbine 120 (also referred to as the core of engine 100) and a rotor assembly (also referred to as fan section 150) positioned upstream thereof. Typically, the turbine 120 includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. 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 shroud 122 at least partially encloses and supports a booster or low pressure ("LP") compressor 126 for pressurizing air entering the turbine 120 through a 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 flow flows downstream to the combustor 130 of the combustion section, where fuel is injected into the pressurized air flow and ignited to raise the temperature and energy level of the pressurized air and produce high energy combustion products.

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 the 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, the low pressure turbine 134 is drivingly coupled with the low pressure compressor 126 and components of the 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 core or 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 shroud 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. As used herein, the term "second flow" or "core flow" refers to flow through the engine inlet and ducted fan and also travels through the core inlet and core duct.

The fan section 150 includes a fan 152, which in this example embodiment is a primary fan. For the embodiment shown in fig. 1, the fan 152 is an open rotor or non-ducted fan 152. As shown, the fan 152 includes an array of fan blades 154 (only one shown in fig. 1). The fan blades 154 are rotatable, for example, about the longitudinal axis 112. As described above, fan 152 is drivingly coupled with low-pressure turbine 134 via LP shaft 138. Fan 152 may be coupled directly with LP shaft 138, for example, in a direct drive configuration. However, for the embodiment shown in FIG. 1, the fan 152 is coupled with the LP shaft 138 via a reduction gearbox 155, for example, 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 vane 154 defines a central vane axis 156. For this embodiment, each fan blade 154 of the fan 152 may rotate about their respective center 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 vary the pitch of the fan blades 154 about their respective center blade axes 156.

The fan section 150 also includes an array of fan guide vanes 160 including 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 unshielded as shown in fig. 1, or alternatively may be unshielded, for example, by an annular shroud spaced outwardly from the tips of the fan guide vanes 162 in 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 may rotate, for example, in unison with each other about their respective center vane axis 164. 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 vane axes 164. However, in other embodiments, each fan guide vane 162 may be fixed or unable to tilt about its central vane axis 164. The fan guide vanes 162 are mounted to a fan case 170.

As shown in FIG. 1, in addition to non-ducted fan 152, ducted fan 184 is included aft of fan 152 such that engine 100 includes both ducted and non-ducted fans that are used to generate thrust by movement of air without passing through at least a portion of turbine 120 (e.g., HP compressor 128 and combustion section for the illustrated embodiment). Ducted fan 184 is shown at approximately the same axial position as fan blades 154 and radially inward of fan blades 154. For the illustrated embodiment, ducted fan 184 is coupled to LP shaft 138 by low-pressure turbine 134 (e.g., coupled to LP shaft)

And (5) driving.

The fan shroud 170 annularly surrounds at least a portion of the core shroud 122 and is positioned generally outboard of at least a portion of the core 5 shroud 122 in the radial direction R. In particular, the downstream section of the fan shroud 170 extends over the front of the core shroud 122,

to define a fan flow path or fan duct 172. The fan flow path or fan duct 172 may be referred to as a third flow of the engine 100. As used herein, the term "third flow" or "fan flow" refers to flow through the engine inlet and ducted fan but not through the core inlet and core duct. Furthermore, the third flow is an air flow that absorbs inlet air, as opposed to free flow air. The third stream passes through at least one stage of the turbine, such as a ducted fan.

0 may enter through fan duct 172 through fan duct inlet 176 and may pass through the fan row

The air jets 178 exhaust to generate propulsion thrust. The fan duct 172 is an annular duct positioned substantially outside the core duct 142 in 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). The stationary struts 174 may each have aerodynamic forces

Contoured to direct the flow of air therethrough. Other struts besides stationary struts 174 may be used to connect and support fan shroud 170 and/or core shroud 122. In many embodiments, the fan duct 172 and the core duct 142 may be at least partially coextensive (generally 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 coextensive generally axially on opposite radial sides of the core cowl.

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 defined generally at the forward end of the fan shroud 170 and is positioned between the fan 152 and the fan guide vane array 160 in the axial direction a. The inlet duct 180 is an annular duct positioned inside the fan housing 170 in the radial direction R. The air flowing downstream along the inlet duct 180 is split by the splitter or leading edge 144 of the core shroud 122 into the core duct 142 and the fan duct 172, not necessarily uniformly. Inlet pipe 180

And is wider in the radial direction R than the core tube 142. The inlet duct 180 is also wider in the radial direction R than the fan duct 172. 5 notably, for the illustrated embodiment, engine 100 includes one or more features to increase third stream thrust

Is not limited to the above-described embodiments. In particular, engine 100 also includes an array of inlet guide vanes ("IGVs") 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 fan inlet guide vanes 186 are not rotatable about the longitudinal axis 112. Each of which

Each inlet guide vane 186 defines a central vane axis (not labeled for clarity) and is rotatable about their respective central 0-center vane axes, e.g., in unison with each other. One or more actuators may be provided to facilitate such rotation and may therefore be used to vary the pitch of the inlet guide vanes 186 about their respective central vane axes. However, in other embodiments, each inlet guide vane 186 may be fixed or unable to tilt 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 ("OGV") 190. As with the array of inlet guide vanes 186, the array of outlet guide vanes 190 is not rotatable about the longitudinal axis 112. However, for the illustrated 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 illustrated 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 may include one or more actuators 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 rate, etc. of the airflow through the fan duct 172). A fixed geometry exhaust nozzle may also be employed.

The combination of the array of inlet guide vanes 186 upstream of the ducted fan 184, the array of outlet guide vanes 190 downstream of the ducted fan 184, and the fan exhaust nozzle 178 may result in more efficient generation of third flow thrust during one or more engine operating conditions. Furthermore, by introducing variability in the geometry of the inlet guide vanes 186 and the fan exhaust nozzle 178, the engine 100 may be able to generate more efficient third flow thrust under a relatively wide range of engine operating conditions, including take-off and climb (typically requiring a maximum total engine thrust) and cruise (typically requiring a smaller amount of total engine thrust).

In exemplary embodiments, the air passing through the fan duct 172 may be relatively cooler (e.g., cooler) than one or more fluids used in the turbine 120. In this manner, one or more heat exchangers 200 may be disposed within the fan duct 172 and used to cool one or more fluids from the turbine 120 (also referred to as the core of the engine 100), wherein air passing through the fan duct 172 acts as a source to remove heat from the fluid (e.g., compressor bleed air, oil, or fuel).

Although not depicted, in certain exemplary embodiments, engine 100 may also include one or more heat exchangers 200 in other annular ducts or flow paths of engine 100 (e.g., in inlet duct 180, in turbine mechanical flow path/core duct 142, within turbine section and/or turbine exhaust nozzle 140, etc.).

In the exemplary embodiment shown, first flow 280 travels outside engine inlet 182. As used herein, the term "first flow" or "free flow" refers to a flow that flows outside the engine inlet and over a fan (which is non-ducted). Further, the first flow 280 is a flow of free-flowing air.

In addition, third flow 284 travels through fan duct 172 and exits fan exhaust nozzle 178. In addition, a second or core stream 282 is also shown along with the first stream 280 and the third stream 284.

Referring now to FIG. 2, a close-up cross-sectional view of a heat transfer system 210 of the present disclosure coupled to or integrated with a portion of the exemplary engine 100 of FIG. 1 is provided.

In such embodiments, the heat transfer system 210 is configured to reduce ice accretion or ice formation in the first engine component (e.g., the inlet guide vanes 186) and the second engine component (e.g., the outlet guide vanes 190).

The heat transfer system 210 is in thermal communication with a heat source. More specifically, for the illustrated embodiment, the heat source is the gearbox 155, and thus the heat transfer system 210 is in communication with the gearbox 155. In this manner, the heat transfer system 210 of the present disclosure utilizes waste heat from the gearbox 155 to reduce ice build-up or ice formation in the components of the engine 100.

However, it should be appreciated that in other embodiments, the heat source may additionally or alternatively be a lubrication oil system of engine 100, an electrical system of engine 100, an airflow discharge of engine 100, optical energy, or any other suitable high temperature system of engine 100.

In the exemplary embodiment, heat transfer system 210 generally includes a gearbox waste heat line 220, a first heat transfer member, and a second heat transfer member. For the illustrated embodiment, the first heat transfer member is a horizontal heat transfer member 230 and the second heat transfer member is a vertical heat transfer member 240. As shown in the depicted embodiment, the heat transfer system 210 receives waste heat from the gearbox 155 via a gearbox waste heat line 220.

As used herein, the term "horizontal heat transfer member" refers to a heat transfer member that extends substantially in a horizontal direction (e.g., parallel to axial direction a) when engine 100 is in a normal operating attitude (e.g., when an aircraft incorporating engine 100 is parked, or engine 100 is in a cruise operating condition). Similarly, the term "vertical heat transfer member" refers to a heat transfer member that extends substantially in a vertical direction (e.g., parallel to the vertical direction along radial direction R) when engine 100 is in a normal operating attitude (e.g., when an aircraft incorporating engine 100 is parked, or engine 100 is in a cruise operating condition).

In the exemplary embodiment, horizontal heat transfer member 230 extends in axial direction A relative to a first engine component (e.g., inlet guide vanes 186) and a second engine component (e.g., outlet guide vanes 190). It is contemplated that horizontal heat transfer member 230 includes a heat pipe configured to transfer heat from gearbox 155 to vertical heat transfer member 240. In this manner, it is understood that horizontal heat transfer members 230 may generally include both a hot interface and a cold interface, and may include phase change material therein. At the thermal interface, the liquid phase change material is in thermal contact with a thermally conductive surface of a heat source (e.g., gearbox 155 in the illustrated embodiment) via a gearbox waste heat line 220, and absorbs heat from the surface and becomes steam. The vapor then travels along the horizontal heat transfer member 230 to the cold interface and condenses back to a liquid, releasing the latent heat of the phase change material. The liquid then returns to the thermal interface by, for example, capillary action or gravity and the cycle repeats. Advantageously, the heat pipe has a high thermal conductivity, e.g. a higher thermal conductivity than a conventional copper rod, due to the efficient heat transfer during the phase transition of the inner material. In other exemplary embodiments, it is contemplated that the heat pipes of the heat transfer system of the present disclosure may have other orientations, such as horizontal, vertical, or tilted, depending on the location of the source and the heat sink.

Moreover, in the exemplary embodiment, vertical heat transfer component 240 extends in a vertical direction along radial direction R within each of the first engine component (e.g., inlet guide vanes 186) and the second engine component (e.g., outlet guide vanes 190). The vertical heat transfer member 240 communicates with the gear case waste heat line 220 via the horizontal heat transfer member 230. It is contemplated that the vertical heat transfer component 240 includes graphene rods or layers configured to transfer heat from the gearbox 155 to the first engine component (e.g., the inlet guide vanes 186) and the second engine component (e.g., the outlet guide vanes 190). In this manner, waste heat from the gearbox 155 is used to reduce ice build-up or ice formation in the first engine component (e.g., the inlet guide vanes 186), the second engine component (e.g., the outlet guide vanes 190), and/or any other desired engine component (e.g., the stationary struts 174). Advantageously, graphene has unique physical properties suitable for the heat transfer systems and anti-icing applications described herein. Graphene has high electrical conductivity and high thermal conductivity, resulting in extremely small heat resistance. Graphene is flexible, thin, and transparent, and can be made into an effective coating that reduces the desired weight. In addition, graphene is stronger, durable, and reduces erosion than steel. Graphene may be applied as a sheet, coating, or as part of an alloy. It is contemplated that graphene rods may replace heat pipes in configurations having multiple bends, space limitations, or working against gravity.

In addition, graphene is flexible, impermeable to molecules, and has high electrical and thermal conductivity. In addition, graphene combines the strength and lightweight properties of carbon network allotropes.

Graphene has a melting temperature of about 5000K (about 4727 ℃) and has desirable properties of being refractory. The conductivity of graphene is anisotropic, and graphene can be used as an insulating material. Graphene also has desirable impact resistance. For these reasons, conventional anti-icing coatings having lower impact resistance are less desirable in aircraft engines.

Referring now to FIG. 3, a close-up cross-sectional view of a heat transfer system 210A according to another exemplary embodiment of the present disclosure is provided, the heat transfer system 210A being coupled to or integrated with a portion of the exemplary engine 100 of FIG. 1.

The exemplary heat transfer system 210A depicted in fig. 3 may be configured in substantially the same manner as the exemplary heat transfer system 210 described above with reference to fig. 2. The embodiment shown in fig. 3 includes similar components to the embodiment shown in fig. 2, and like components are denoted by the reference numeral followed by the letter a. For brevity, these similar components of the heat transfer system 210A (FIG. 3) will not all be discussed in connection with the embodiment shown in FIG. 3.

In the exemplary embodiment shown, heat transfer system 210A also includes a second horizontal heat transfer member 250 that extends in an axial direction A from gearbox 155 toward forward end 114 of engine 100. In this manner, the heat transfer system 210A is configured to provide waste heat from the gearbox 155 to reduce ice build-up or ice formation in the front and rear components of the gearbox 155 of the engine 100. For example, the second horizontal heat transfer member 250 extends forward of the gearbox 155 and the horizontal heat transfer member 230A extends rearward of the gearbox 155.

Moreover, in the exemplary embodiment shown, horizontal heat transfer member 230A may extend in axial direction A between the first engine component and the second engine component at a location radially inward from inlet duct 180. In the exemplary embodiment shown in FIG. 2, horizontal heat transfer member 230 extends in axial direction A between the first engine component and the second engine component at a location radially outward from inlet duct 180. In other exemplary embodiments, it is contemplated that the heat transfer system of the present disclosure may include horizontal graphene rods connected to vertical heat pipes, and vice versa.

Referring now to FIG. 4, a close-up cross-sectional view of a heat transfer system 210B according to another exemplary embodiment of the present disclosure is provided, the heat transfer system 210B being coupled to or integrated with a portion of the exemplary engine 100 of FIG. 1.

The exemplary heat transfer system 210B depicted in fig. 4 may be configured in substantially the same manner as the exemplary heat transfer system 210 described above with reference to fig. 2. The embodiment shown in fig. 4 includes similar components to the embodiment shown in fig. 2, and like components are denoted by the reference numerals followed by the letter B. For brevity, these similar components of the heat transfer system 210B (fig. 4) will not all be discussed in connection with the embodiment shown in fig. 4.

In the exemplary embodiment shown, heat transfer system 210B uses a vertical heat transfer component 240B, with vertical heat transfer component 240B extending in a radial direction R within a first engine component (e.g., inlet guide vanes 186) and being in direct thermal communication with a portion of a gearbox (e.g., gearbox hot oil pipe 290). It is contemplated that the gearbox hot oil pipe 290 contains lubrication oil from the gearbox 155. Such a configuration enables the inlet guide vanes 186 and the splitter portion, such as the splitter or leading edge 144 (fig. 1 and 2), to be anti-icing circumferentially with a relatively small amount of oil.

FIG. 5 provides a block diagram of an exemplary control system 400 for controlling engine 100 (FIG. 1) and any heat transfer systems 210, 210A, 210B (FIGS. 2-4) in accordance with an exemplary embodiment of the present disclosure.

Referring to fig. 5, a control system 400 of the present disclosure may be in communication with heat transfer systems 210, 210A, 210B (fig. 2-4) of engine 10. For example, the control system 400 may be used to determine when to activate the heat transfer systems 210, 210A, 210B (fig. 2-4) of the present disclosure to provide heat from the gearbox 155 to selected engine components via the heat transfer systems 210, 210A, 210B (fig. 2-4).

In some embodiments, all of the components of control system 400 are on turbofan engine 10. In other embodiments, some components of control system 400 are on turbofan engine 100, while some are external to turbofan engine 100. For example, some of the external components may be mounted to a wing, fuselage, or other suitable structure of an aircraft in which turbofan engine 100 is mounted.

Still referring to fig. 5, the control system 400 includes a controller 410 and a sensing unit 420 (e.g., including temperature sensors (e.g., T12, T3, T25), or other temperature sensors, pressure sensors (e.g., P3 or other pressure sensors), and/or speed sensors), and is operable with an engine having a heat source 430 and heat transfer systems 210, 210A, 210B. In the exemplary embodiment, control system 400 and heat transfer systems 210, 210A, 210B are in communication with a first set of engine components 440 and a second set of engine components 450. It is contemplated that the first engine component (e.g., inlet guide vanes 186 (FIG. 2)) may be one or more of first set of engine components 440 and the second engine component (e.g., outlet guide vanes 190 (FIG. 2)) may be one or more of second set of engine components 450.

It is also contemplated that first set of engine components 440 may include a splitter component, a booster inlet guide vane, a rotating cone component, and/or other similar engine components. Further, it is contemplated that second set of engine components 450 may include outlet guide vanes, strut components, temperature sensor components, high pressure compressor inlet guide vanes, variable bleed valve ("VBV") components, and/or other similar engine components.

In an exemplary embodiment, the sensing unit 420 may include sensors located at desired components of the engine 10. The sensing unit 420 of the control system 400 monitors the conditions 455 of the components of the engine 100. In general, when sensing unit 420 receives an input indicating a change in a condition 455 of one of the components of engine 100, controller 410 may determine one or more conditions indicating an icing condition or potential icing condition and, at activation of heat pipe/graphene rod loop control 490, activate a loop of heat transfer system 210 of the present disclosure to provide heat from heat source 430 to a detected component of engine 100 via heat transfer system 210. The circuit may be a heat pipe/graphene rod circuit control 490 and the heat transfer system 210 may include a plurality of circuits extending in thermal communication with a plurality of different locations, including the first set of engine components 440 and the second set of engine components 450.

It is contemplated that conditions 455 of components of engine 100 monitored by sensing unit 420 include temperature, pressure, and/or other information indicative of icing conditions and/or ice formation on components of engine 100.

Advantageously, the control system 400 of the present disclosure may monitor the first type of detection information 460 and the second type of detection information 470. In this manner, in response to receiving an input type indicative of a change in a condition of a component of engine 100, control system 400 is configured to selectively provide heat from heat source 430 to first set of engine components 440 or second set of engine components 450.

It is contemplated that the first type of detection information 460 includes supercooled liquid (SCL) information. For example, such information may include sensing some water droplets at a predetermined lower temperature. Further, the first type detection information 460 may correspond to the first set of engine components 440. Supercooled liquid (SCL) information is the state of water droplets below freezing, which tend to form ice when striking aircraft/engine surfaces.

It is also contemplated that second type of detection information 470 includes Ice Crystal Ice (ICI) detection information through a flow path of engine 100 and may include temperature information. Further, the second type of detection information 470 may correspond to a second set of engine components 450. Ice Crystal Ice (ICI) information is a solid state of high altitude water droplets that tend to bounce, chip, melt and possibly adhere to aircraft/engine surfaces.

The first set of engine components 440 and the second set of engine components 450 are susceptible to icing under different conditions. Thus, having the first type of detection information 460 for the first set of engine components 440 and the second type of detection information 470 for the second set of engine components 450 allows for proper icing monitoring and detection for each set of engine components. Further, in this manner, the control system 400 can activate the loop of the heat transfer system 210 for only the set of engine components experiencing a potential icing condition, for example, at the activation heat pipe/graphene rod loop control 490. For example, the active heat pipe/graphene rod loop control 490 of the present disclosure allows for selective activation of only loops to the first set of engine components 440 or the second set of engine components 450. In this manner, control system 400 is able to selectively activate and selectively provide heat from heat source 430 to only detected components of engine 100 via heat transfer system 210. For example, in the exemplary embodiment, active heat pipe/graphene rod loop control 490 includes a mechanical switch that is actuated by a Full Authority Digital Engine Control (FADEC) system to complete a loop between a source and a radiator to transfer heat, such as a connector between vertical heat transfer component 240B (fig. 4) and hot oil pipe 290 (fig. 4).

In addition, control system 400 includes a pilot command portion 480. In the exemplary embodiment, pilot command portion 480 indicates an operating condition in which engine 100 (FIG. 2) is susceptible to icing or ice formation.

In the exemplary embodiment, engine 100 (FIG. 2) also includes a computing system. Specifically, for this embodiment, engine 100 includes a computing system having one or more computing devices, including controller 410, controller 410 being configured to control engine 100, and in this embodiment, heat source 430 and other components of control system 400. The controller 410 may include one or more processors and associated memory devices configured to perform various computer-implemented functions and/or instructions (e.g., perform methods, steps, calculations, etc., and store relevant data as disclosed herein). The instructions, when executed by the one or more processors, may cause the one or more processors to operate, for example, to provide heat from the gearbox 155 to the engine components via the heat transfer system 210 upon receiving an input indicative of a change in a condition of one of the components of the engine 100.

Further, controller 410 may include a communication module to facilitate communication between controller 410 and various components of the aircraft and other electrical components of engine 100. The communication module may include a sensor interface (e.g., one or more analog-to-digital converters) to allow signals transmitted from the one or more sensors to be converted into signals that can be understood and processed by the one or more processors. It should be appreciated that the sensor may be communicatively coupled to the communication module using any suitable means. For example, the sensor may be coupled to the sensor interface via a wired connection. However, in other embodiments, the sensor may be coupled to the sensor interface via a wireless connection (e.g., by using any suitable wireless communication protocol). Thus, the processor may be configured to receive one or more signals or outputs, such as one or more operating conditions/parameters, from the sensors.

As used herein, the term "processor" refers not only to integrated circuits referred to in the art as being included in a computing device, but also to controllers, microcontrollers, microcomputers, programmable Logic Controllers (PLCs), application specific integrated circuits, and other programmable circuits. The one or more processors may also be configured to perform the computations necessary to execute the advanced algorithms. Furthermore, the memory device 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), floppy disks, compact disk read-only memories (CD-ROMs), magneto-optical disks (MODs), digital Versatile Disks (DVDs), and/or other suitable memory elements. Such memory devices may generally be configured to store suitable computer readable instructions that, when implemented by a processor, configure the controller 410 to perform the various functions described herein. The controller 410 may be configured in substantially the same manner as the exemplary computing device of the computing system 500 described below with reference to fig. 7 (and may be configured to perform one or more functions of the exemplary method (550) described herein).

The controller 410 may be a controller system or a single controller. Controller 410 may be a controller dedicated to controlling heat source 430 and associated electrical components, or may be an engine controller configured to control engine 100 and control system 400 and its associated electrical components. The controller 410 may be an Electronic Engine Controller (EEC) or Electronic Control Unit (ECU), such as a Full Authority Digital Engine Control (FADEC) system.

The control system 400 may include one or more power management electronics or electrical control devices, such as inverters, converters, rectifiers, devices operable to control the flow of current, and the like. For example, one or more of the control devices may be operable to regulate and/or convert power (e.g., from AC to DC, or vice versa). Further, one or more of the control devices may be operable to control the amount of heat provided to engine components by heat source 430 via heat transfer system 210. Although the control devices may be separate from the heat source 430 and the controller 410, it should be understood that one, some, or all of the control devices may be located on the heat source 430 and/or the controller 410.

As discussed, engine 100 may also include one or more sensors for sensing and/or monitoring various engine operating conditions and/or parameters during operation. The sensors of the sensing unit 420 may sense or measure various engine conditions (e.g., pressure and temperature), and one or more signals may be directed from the one or more sensors to the controller 410 for processing. Accordingly, the controller 410 is communicatively coupled with one or more sensors, for example, via a suitable wired or wireless communication link. It should be appreciated that engine 100 may include other sensors located at other suitable locations along the core air flow path.

In an exemplary embodiment, one or more sensors of sensing unit 420 may monitor the temperature of engine 100, and controller 410 may be configured to activate heat source 430 once certain predetermined conditions of components of engine 100 have been reached. In an exemplary embodiment, the one or more sensors of the sensing unit 420 may include a temperature detector.

FIG. 6 provides a flowchart of an exemplary method (550) of providing heat from a gearbox 155 of an engine 100 to a first engine component and a second engine component of the engine 100 to reduce ice accretion or ice formation in accordance with an exemplary embodiment of the present disclosure. For example, exemplary method (550) may be used to operate engine 100 as described herein. It should be understood that the method (550) discussed herein is merely to describe exemplary aspects of the present subject matter and is not intended to be limiting.

At (552), the method (550) includes receiving, by one or more computing devices, an input indicative of a change in a condition of a portion of the turbofan engine 100. For example, controller 410 may receive input in response to when a condition (e.g., temperature or pressure) of a component of engine 100 is reached.

At 554, responsive to receiving the type of input indicative of a change in the condition of the portion of the engine 100, the method 550 includes selectively providing, by the one or more computing devices, heat from the gearbox 155 to one of the first and second engine components.

FIG. 7 provides an example computing system 500 according to an example embodiment of the disclosure. For example, the computing system described herein (e.g., controller 410) may include various components and perform the various functions of computing system 500 described below.

As shown in fig. 7, computing system 500 may include one or more computing devices 510. Computing device 510 may include one or more processors 510A and one or more memory devices 510B. The one or more processors 510A may include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, and/or other suitable processing device. The one or more memory devices 510B may include one or more computer-readable media including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard disk drives, flash drives, and/or other memory devices.

The one or more memory devices 510B may store information accessible by the one or more processors 510A, including computer-readable instructions 510C that may be executed by the one or more processors 510A. The instructions 510C may be any set of instructions that, when executed by the one or more processors 510A, cause the one or more processors 510A to perform operations. In some embodiments, the instructions 510C may be executable by the one or more processors 510A to cause the one or more processors 510A to perform operations, such as any operations and functions for which the computing system 500 and/or the computing device 510 is configured, for electrically assisting operation of the turbine during transient operations (e.g., the method (550) of fig. 6), and/or any other operations or functions of the one or more computing devices 510. Thus, the method (550) may be a computer-implemented method such that each step of the exemplary method (550) is performed by one or more computing devices (e.g., the exemplary computing device 510 of the computing system 500). The instructions 510C may be software written in any suitable programming language or may be implemented in hardware. Additionally and/or alternatively, the instructions 510C may execute in logically and/or virtually separate threads on the one or more processors 510A. The one or more memory devices 510B may also store data 510D that is accessible by the one or more processors 510A. For example, data 510D may include models, databases, and the like.

Computing device 510 may also include a network interface 510E for communicating with other components of system 500 (e.g., via a network), for example. Network interface 510E may include any suitable components for interfacing with one or more networks, including, for example, a transmitter, a receiver, a port, a controller, an antenna, and/or other suitable components. One or more external devices (e.g., electrical control devices) may be configured to receive one or more commands from computing device 510 or to provide one or more commands to computing device 510.

The control system 400 of the present disclosure does not require modification of the mechanical hardware of the engine and facilitates simple retrofitting of existing engines.

It is contemplated that the turbines and methods of the present disclosure may be implemented on an aircraft, helicopter, automobile, ship, submarine, train, unmanned aerial vehicle, or drone, and/or any other suitable vehicle. Although the present disclosure is described herein with reference to aircraft embodiments, this is intended to be by way of example only and not by way of limitation. Those of ordinary skill in the art will appreciate that the turbines and methods of the present disclosure may be implemented on other vehicles without departing from the scope of the present disclosure.

The techniques discussed herein refer to computer-based systems, actions taken by computer-based systems, information sent to computer-based systems, and information from computer-based systems. Those of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for a variety of possible configurations, combinations, and divisions of tasks and functions between and among components. For example, the processes discussed herein may be implemented using a single computing device or multiple computing devices working in combination. The databases, memories, instructions and applications may be implemented on a single system or distributed across multiple systems. The distributed components may operate sequentially or in parallel.

Although specific features of various embodiments may be shown in some drawings and not in others, this is for convenience only. Any feature of the drawings may be referenced and/or claimed in combination with any feature of any other drawing in accordance with the principles of the present disclosure.

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

a turbofan engine defining an axial direction and a radial direction, the turbofan engine comprising: a fan; a turbine operatively coupled to the fan to drive the fan, wherein the turbine, the fan, or both comprise an engine component; a heat source; and a heat transfer system configured to reduce ice accumulation or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising: a first heat transfer member in communication with the heat source; and a second heat transfer member extending from the first heat transfer member to or through the engine member, wherein the first heat transfer member comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer member comprises the other of the heat pipe or the graphene rod.

A turbofan engine according to any preceding claim, wherein the first heat transfer member is a horizontal heat transfer member, and wherein the second heat transfer member is a vertical heat transfer member.

A turbofan engine according to any preceding claim, wherein the horizontal heat transfer component comprises the heat pipe configured to transfer the heat from the heat source to the vertical heat transfer component.

The turbofan engine of any preceding clause, wherein the vertical heat transfer component comprises the graphene rods configured to transfer the heat to the engine component.

A turbofan engine according to any preceding claim, wherein the heat source is a gearbox, and wherein the turbine is operatively coupled to the fan through the gearbox.

A turbofan engine according to any preceding clause, wherein the engine component is a first engine component, wherein the turbine, the fan, or both further comprise a second engine component, and wherein the heat transfer system is configured to selectively transfer heat from the gearbox to the first and second engine components.

The turbofan engine of any preceding clause, further comprising a controller having one or more processors and one or more memory devices storing instructions that, when executed by the one or more processors, cause the one or more processors to perform operations that, when performed, are configured to: receiving a first input indicative of a first change in a condition of a first portion of the turbofan engine; and providing the heat from the heat source to the first engine component via the heat transfer system in response to the first input.

The turbofan engine of any preceding clause, wherein the one or more processors are further configured to: receiving a second input indicative of a second change in a condition of a second portion of the turbofan engine; and providing the heat from the heat source to the second engine component via the heat transfer system in response to the second input.

A turbofan engine according to any preceding clause, wherein the first engine component is one of a splitter component, a booster inlet guide vane component, and a rotating cone component.

A turbofan engine according to any preceding clause, wherein the second engine component is one of an outlet guide vane component, a strut component and a compressor inlet guide vane component.

The turbofan engine of any preceding clause wherein the fan is a non-ducted fan.

A heat transfer system for a turbofan engine defining an axial direction and a radial direction, the turbofan engine comprising: a fan; a turbine operatively coupled to the fan to drive the fan, wherein the turbine, the fan, or both comprise an engine component; and a heat source, the heat transfer system configured to reduce ice accumulation or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising: a first heat transfer member in communication with the heat source; and a second heat transfer member extending from the first heat transfer member to or through the engine member, wherein the first heat transfer member comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer member comprises the other of the heat pipe or the graphene rod.

The heat transfer system of any preceding claim, wherein the first heat transfer member is a horizontal heat transfer member, and wherein the second heat transfer member is a vertical heat transfer member.

The heat transfer system of any preceding clause, wherein the horizontal heat transfer component comprises the heat pipe configured to transfer the heat from the heat source to the vertical heat transfer component, and wherein the vertical heat transfer component comprises the graphene rod configured to transfer the heat to the first and second engine components.

The heat transfer system of any preceding clause, wherein the heat source is a gearbox, and wherein the turbine is operably coupled to the fan through the gearbox.

The heat transfer system of any preceding clause, wherein the engine component is a first engine component, wherein the turbine, the fan, or both further comprise a second engine component, and wherein the heat transfer system is configured to selectively transfer heat from the gearbox to the first and second engine components.

A method of providing heat from a gearbox of a turbofan engine to first and second engine components of the turbofan engine to reduce ice build-up or ice formation, the turbofan engine defining an axial direction and a radial direction and including a fan and a turbine operatively coupled to the fan to drive the fan, the method comprising: receiving, by one or more computing devices, an input indicative of a change in a condition of a portion of the turbofan engine; and selectively providing, by the one or more computing devices, the heat from the gearbox to one of the first and second engine components in response to a type of received input indicative of the change in the condition of the portion of the turbofan engine.

The method of any preceding clause, further comprising: receiving, by the one or more computing devices, a first input indicative of a first change in a condition of a first portion of the turbofan engine; and providing, by the one or more computing devices, the heat from the gearbox to the first engine component in response to the received first input indicative of the first change in the condition of the first portion of the turbofan engine.

The method of any preceding clause, further comprising: receiving, by the one or more computing devices, a second input indicative of a second change in a condition of a second portion of the turbofan engine; and providing, by the one or more computing devices, the heat from the gearbox to the second engine component in response to the received second input indicative of the second change in the condition of the second portion of the turbofan engine.

The method of any preceding claim, wherein selectively providing, by the one or more computing devices, the heat from the gearbox to one of the first and second engine components comprises providing the heat to the one of the first and second engine components in the axial direction, and providing the heat within the one of the first and second engine components in the radial direction.

A turbofan engine according to any preceding claim, wherein the second horizontal heat transfer member extends forward of the gearbox and the horizontal heat transfer member extends aft of the gearbox.

A turbofan engine according to any preceding claim, wherein the horizontal heat transfer member extends in the axial direction between the first and second engine members at a position radially inward from the inlet duct 5.

A turbofan engine according to any preceding claim, wherein the horizontal heat transfer member extends in the axial direction between the first and second engine members at a location radially outward from the inlet duct.

A turbofan engine according to any preceding claim, wherein the heat transfer system comprises a vertical heat transfer portion

The vertical heat transfer member extends in said radial direction within the inlet guide vanes and is in direct thermal communication with a portion of the gearbox hot oil pipe 0.

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. If these other

Examples 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 5 language of the claims, then such other examples are intended to be within the scope 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.

Claims (10)

1. A turbofan engine defining an axial direction and a radial direction, the turbofan engine comprising:

a fan;

a turbine operatively coupled to the fan to drive the fan, wherein the turbine, the fan, or both comprise an engine component;

a heat source; and

a heat transfer system configured to reduce ice accumulation or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising:

A first heat transfer member in communication with the heat source; and

a second heat transfer member extending from the first heat transfer member to or through the engine member, wherein the first heat transfer member comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer member comprises the other of the heat pipe or the graphene rod.

2. The turbofan engine of claim 1 wherein the first heat transfer member is a horizontal heat transfer member and wherein the second heat transfer member is a vertical heat transfer member.

3. The turbofan engine of claim 2 wherein the horizontal heat transfer component comprises the heat pipe configured to transfer the heat from the heat source to the vertical heat transfer component.

4. The turbofan engine of claim 2 wherein the vertical heat transfer component comprises the graphene rod configured to transfer the heat to the engine component.

5. The turbofan engine of claim 1 wherein the heat source is a gearbox and wherein the turbine is operatively coupled to the fan through the gearbox.

6. The turbofan engine of claim 5 wherein the engine component is a first engine component, wherein the turbine, the fan, or both further comprise a second engine component, and wherein the heat transfer system is configured to selectively transfer heat from the gearbox to the first engine component and the second engine component.

7. The turbofan engine of claim 6 further comprising

A controller having one or more processors and one or more memory devices storing instructions that when executed by the one or more processors cause the one or more processors to perform operations that, when performed, are configured to:

receiving a first input indicative of a first change in a condition of a first portion of the turbofan engine; and is also provided with

In response to the first input, the heat is provided from the heat source to the first engine component via the heat transfer system.

8. The turbofan engine of claim 7 wherein the one or more processors are further configured to:

receiving a second input indicative of a second change in a condition of a second portion of the turbofan engine; and is also provided with

In response to the second input, the heat is provided from the heat source to the second engine component via the heat transfer system.

9. The turbofan engine of claim 6 wherein the first engine component is one of a splitter component, a booster inlet guide vane component, and a rotating cone component.

10. The turbofan engine of claim 9 wherein the second engine component is one of an outlet guide vane component, a strut component, and a compressor inlet guide vane component.