阅读说明：本技术 具有热管理系统的发动机组件 (Engine assembly with thermal management system ) 是由 S.贝格龙 E.普拉蒙顿 A.朱利安 于 2020-02-03 设计创作，主要内容包括：公开了一种用于飞行器的发动机组件,其包括：燃烧发动机,其包括与散热器处于热交换关系的冷却剂回路,该散热器包括热交换器和至少一个部件,该至少一个部件具有不同于热交换的主要功能。还公开了一种操作系统的方法。(An engine assembly for an aircraft is disclosed, comprising: a combustion engine comprising a coolant circuit in heat exchange relationship with a radiator, which radiator comprises a heat exchanger and at least one component having a primary function other than heat exchange. A method of operating a system is also disclosed.)

1. An engine assembly for an aircraft, comprising: a combustion engine comprising a coolant circuit in heat exchange relationship with a radiator, the radiator comprising a heat exchanger and at least one further component of the engine assembly, the at least one further component having a primary function other than heat exchange.

2. The engine assembly of claim 1, wherein the coolant circuit includes a conduit in contact with the at least one additional component, the conduit being thermally conductively connected with the at least one additional component.

3. The engine assembly of claim 1, wherein the coolant circuit includes a conduit in heat exchange relationship with the at least one additional component, the conduit being selectively connectable to the remainder of the coolant circuit.

4. The engine assembly of claim 1, wherein the at least one additional component is selected from the group consisting of: a gearbox, an inlet, a compressor, a turbine, a generator, an engine starter, a load compressor, an air filter, an actuator, and a valve.

5. The engine assembly of claim 1, wherein the at least one additional component is operatively connected to the combustion engine.

6. The engine assembly of claim 1, further comprising a heat source in heat exchange relationship with the coolant circuit and with the at least one additional component.

7. The engine assembly of claim 1, wherein the coolant circuit includes a pump drivingly engaged with an engine shaft of the combustion engine to induce the flow of the liquid coolant in the coolant circuit.

8. The engine assembly of claim 1, wherein the coolant circuit includes a pump selectively drivingly engaged by an electric motor operatively connected to a power source or by an engine shaft of the combustion engine.

9. The engine assembly of claim 8, further comprising a clutch having an input drivingly engaged by an engine shaft and an output drivingly engaged with the pump, the clutch operable in a first configuration in which the input is drivingly engaged with the output and a second configuration in which the input is disengaged from the output.

10. The engine assembly of claim 1, said combustion engine being a reciprocating engine.

11. A method of operating a thermal management system for an engine assembly of an aircraft, comprising:

circulating a liquid coolant;

transferring heat from a combustion engine of the engine assembly to a liquid coolant; and

transferring heat from the liquid coolant to an environment external to the engine assembly via both a heat exchanger and components whose primary function is different from heat exchange.

12. The method of claim 11, wherein transferring heat to the environment via the component comprises transferring heat to a gearbox of the engine assembly.

13. The method of claim 11, wherein transferring heat to the environment via the component comprises transferring heat to the component operably connected to the combustion engine.

14. The method of claim 11, wherein circulating the liquid coolant includes actuating a pump with one of an electric motor and an engine shaft of the combustion engine.

15. The method of claim 11, actuating the pump with the electric motor includes disengaging the engine shaft from the pump.

16. A method of operating a thermal management system for an engine assembly of an aircraft, the engine assembly comprising a combustion engine cooled by a liquid coolant liquid, the method comprising:

determining that the combustion engine of the engine assembly is off;

circulating the liquid coolant in a coolant loop for heating the liquid coolant;

and

circulating heated liquid coolant toward at least one component whose primary function is different from heat exchange for transferring heat from the heated liquid coolant to the at least one component.

17. The method of claim 16, wherein determining that the combustion engine is off includes determining that the combustion engine is cold, and heating the liquid coolant includes heating the liquid coolant with a heat source.

18. The method of claim 16, wherein determining that the combustion engine is off includes determining that the combustion engine is hot, and heating the liquid coolant includes transferring heat from the combustion engine to the liquid coolant.

19. The method of claim 16, wherein determining that the combustion engine of the engine assembly is off further comprises determining that the engine assembly is below a given temperature.

20. The method of claim 16, wherein determining that the combustion engine of the engine assembly is off further comprises determining that a flight phase of the aircraft is an approaching phase.

Technical Field

The present application relates generally to aircraft engines and more particularly relates to systems and methods for managing heat generated by such engines.

Background

Aircraft engines do not perform as well during start-up as during steady state operation because the various engine components and fluids are not at their optimal operating temperatures. After an aircraft engine is shut down, a significant amount of time passes before the engine cools. Furthermore, it has been a long-standing challenge to provide adequate cooling to the hot components of aircraft engines. Therefore, improvements are suitable.

Disclosure of Invention

In one aspect, there is provided an engine assembly for an aircraft, comprising: a combustion engine comprising a coolant circuit in heat exchange relationship with a radiator, the radiator comprising a heat exchanger and at least one component having a primary function different from heat exchange.

In another aspect, a method of operating a thermal management system for an engine assembly of an aircraft is provided, the method comprising: circulating a liquid coolant; transferring heat from a combustion engine of the engine assembly to the liquid coolant; and transferring heat from the liquid coolant to an environment external to the engine assembly via both the heat exchanger and components having primary functions other than heat exchange.

In yet another aspect, a method of operating a thermal management system for an engine assembly of an aircraft, the engine assembly comprising a combustion engine liquid-cooled with a liquid coolant, the method comprising: determining that a combustion engine of the engine assembly is off; circulating a liquid coolant in a coolant loop for heating the liquid coolant; and circulating the heated liquid coolant toward at least one component whose primary function is different from heat exchange for transferring heat from the heated liquid coolant to the at least one component.

Drawings

Referring now to the drawings wherein:

FIG. 1 is a schematic cross-sectional view of a rotating internal combustion engine according to certain embodiments;

FIG. 2 is a schematic view of an engine assembly according to one embodiment; and

FIG. 3 is a schematic illustration of a control system for the engine assembly of FIG. 2.

Detailed Description

Referring to fig. 1, a rotary internal combustion engine 10, referred to as a wankel engine, is schematically illustrated. The rotary engine 10 includes an outer body 12 having axially spaced end walls 14 with a peripheral wall 18 extending between the end walls 14 to form a rotor cavity 20. The inner surface of the peripheral wall 18 of the cavity 20 has a profile defining two lobes, which is preferably a epitrochoid.

An inner body or rotor 24 is received within the cavity 20. The rotor 24 has axially spaced end faces 26 adjacent the outer body end wall 14 and a peripheral face 28 extending therebetween. The peripheral surface 28 defines three circumferentially spaced apart apex portions 30 and a generally triangular profile having outwardly arcuate sides 36. The apex portions 30 sealingly engage the inner surface of the peripheral wall 18 to form three rotating combustion chambers 32 between the inner rotor 24 and the outer body 12. The geometric axis of the rotor 24 is offset from and parallel to the axis of the outer body 12.

The combustion chamber 32 is sealed. In the illustrated embodiment, each rotor apex portion 30 has an apex seal 52, the apex seal 52 extending from one end face 26 to the other and being biased radially outwardly against the peripheral wall 18. An end seal 54 engages each end of each apex seal 52 and is biased against the respective end wall 14. Each end face 26 of the rotor 24 has at least one arcuate face seal 60, the arcuate face seal 60 extending from each apex portion 30 to each adjacent apex portion 30, adjacent to but within the periphery of the rotor throughout its length, in sealing engagement with the end seals 54 adjacent each end thereof, and biased into sealing engagement with the adjacent end wall 14. Alternative sealing arrangements are also possible.

Although not shown in the drawings, the rotor 24 is journalled on an eccentric portion of the shaft such that the shaft rotates the rotor 24 to perform an orbital revolution within the stator cavity 20. As the rotor 24 moves around the stator cavity 20, the shaft rotates 3 times for each complete rotation of the rotor 24. An oil seal is provided around the eccentric to prevent leakage flow of lubricating oil radially outwardly between the respective rotor end face 26 and the outer body end wall 14. During each rotation of the rotor 24, each chamber 32 changes in volume and moves around the stator cavity 20 to undergo four phases of intake, compression, expansion and exhaust, similar to strokes in a reciprocating internal combustion engine with a four stroke cycle.

The engine includes a primary inlet port 40 in communication with an air source, an exhaust port 44, and an optional purge port 42 also in communication with the air source (e.g., a compressor) and located between the inlet port 40 and the exhaust port 44. The ports 40, 42, 44 may be defined in the end wall 14 in the peripheral wall 18. In the illustrated embodiment, an inlet port 40 and a purge port 42 are defined in the end wall 14 and communicate with the same intake conduit 34, the intake conduit 34 being defined as a passage in the end wall 14, and an exhaust port 44 being defined through the peripheral wall 18. Alternative configurations are possible.

In particular embodiments, a fuel such as kerosene (jet fuel) or other suitable fuel is delivered into the chamber 32 through a fuel port (not shown) such that the chamber 32 is stratified with a rich fuel-air mixture near the ignition source and a leaner mixture elsewhere, and the fuel-air mixture may be ignited within the housing using any suitable ignition system known in the art (e.g., spark plug, glow plug). In a particular embodiment, rotary engine 10 operates under the principles of the Miller or Atkinson cycle with a compression ratio lower than its expansion ratio through appropriate relative positioning of primary inlet port 40 and exhaust port 44.

Referring now to FIG. 2, an engine assembly is shown generally at 100. the engine assembly 100 may be an auxiliary power unit for an aircraft the engine assembly 100 includes an internal combustion engine 110, which internal combustion engine 110 may be the rotary engine 10 described above with reference to FIG. 1. alternatively, the internal combustion engine 110 may be any suitable engine the internal combustion engine 110 may be a reciprocating engine, such as a piston engine in certain embodiments, the internal combustion engine 110 may be part of an engine system that is a compound cycle engine system or a compound cycle engine, such as described in U.S. Pat. No. 7,753,036 to L ents et al published on 7/13/2010, or such as described in U.S. Pat. No. 7,775,044 to Julie et al published on 8/17/2010, or such as described in U.S. Pat. publication No. 2015/0275749 to Thomasin et al published on 10/1/2015, or such as U.S. Pat. publication No. 2015/0275756 to Boldouc et al published on 10/1/2015, the contents of which are incorporated herein by reference in their entirety.

The internal combustion engine 110 is fluidly connected to an engine inlet 111, and the engine inlet 111 is fluidly connected to an environment E external to the engine assembly 100.

In the illustrated embodiment, the engine assembly 100 includes a gearbox 112 in driving engagement with an internal combustion engine 110. More specifically, the internal combustion engine 110 has an engine shaft 110a connected to an input 112a of a gearbox 112. The gearbox 112 may be connected to a plurality of components to transmit rotational input from the engine shaft 110 a. The gearbox 112 may produce a rotational speed ratio between its input 112a and its output 112 b.

In the illustrated embodiment, the engine assembly 100 includes a thermal management system 114. The thermal management system 114 includes a coolant loop 116 configured to circulate a liquid coolant. The coolant loop 116 is in heat exchange relationship with the internal combustion engine 110. In a particular embodiment, the internal combustion engine 110 includes a housing, such as the peripheral wall 18 of the rotary engine of FIG. 1, that defines a conduit therein; the conduits are fluidly connected to a coolant circuit 114. The liquid coolant may be able to cool the internal combustion engine 110 by absorbing heat from the housing of the internal combustion engine 110 via convection. Thus, the liquid coolant may increase in temperature after it passes through the housing.

To induce a flow of liquid coolant within the coolant loop 116, the engine assembly 100 includes a pump 118 fluidly connected to the coolant loop 116. The pump 118 may be drivingly engaged by the internal combustion engine 110 and/or by an electric motor 120 powered by the power source S. In a particular embodiment, the power source S is a battery 122. Alternatively, power source S may be an electrical generator drivingly engaged by internal combustion engine 110 or another engine of an aircraft containing engine assembly 100.

In the illustrated embodiment, the thermal management system 114 includes an expansion tank 117 fluidly connected to the coolant circuit 116. The expansion tank 117 may be pressurized at a given pressure (which may be 35 PSI) and may be used for the disposal of thermally induced volume changes of the liquid coolant.

In some cases, it may be useful to circulate liquid coolant in the coolant loop 116 when the internal combustion engine 110 is powered off, shut down, or shut down. Herein, turning off power, shutting down or off means that the internal combustion engine does not cause rotation of the engine shaft 110 a. In other words, by shutting down power, shutting down, or shutting down, no combustion occurs in the combustion chamber(s) of the internal combustion engine 110.

In the illustrated embodiment, the pump 118 is selectively drivingly engaged by the internal combustion engine 110 or the electric motor 120. In the depicted embodiment, the pump 118 is drivingly engaged by the internal combustion engine via the gearbox 112.

To allow the pump 118 to be engaged with the electric motor 120, the engine assembly 100 includes a clutch 124, the clutch 124 having an input 124a and an output 124b, the input 124a being in driving engagement with the internal combustion engine 110 via the gearbox 112, and the output 124b being in driving engagement with the pump 118. The clutch 124 is operable in a first configuration, wherein the input 124a is drivingly engaged with the output 124b, and the clutch 124 is operable in a second configuration, wherein the input 124a is disengaged from the output 124 b. Thus, in the first configuration, the internal combustion engine 110 drives the pump 118, and in the second configuration, the internal combustion engine 110 is decoupled from the pump 118. By being in the second configuration, the clutch 124 allows the electric motor 120 to drive the pump 118 without having to overcome the load created by the internal combustion engine 110 being shut down.

In a particular embodiment, the clutch 124 is an overrunning clutch (sprag clutch). In an overrunning clutch, if input 124a rotates at a greater speed than output 124b, input 124a drivingly engages output 124b, and if output 124b rotates at a greater speed than input 124a, input 124a allows output 124b to rotate independently of input 124 a.

The thermal management system 114 also includes a heat sink 126 in heat exchange relationship with the coolant loop 116. In the illustrated embodiment, the radiator 126 includes a heat exchanger 128. The heat generated by the engine 110 must be dissipated in the environment E. A heat exchanger 128 may be used for this purpose.

More specifically, and in the illustrated embodiment, the heat exchanger 128 includes at least one first conduit 128a and at least one second conduit 128b in heat exchange relationship with the at least one first conduit 128 b. At least one first conduit 128a is fluidly connected to the coolant circuit 116, and at least one second conduit 128b is fluidly connected to the environment E. Blower B may be used to direct air from ambient E through at least one second conduit 128B of heat exchanger 128. In certain embodiments, the blower B may be replaced by scoops located on the exterior surface of the aircraft. The temperature of the air in the environment E is typically lower than the temperature of the liquid coolant that absorbs heat from the internal combustion engine 110. Thus, within the heat exchanger 128, heat transfer occurs from the liquid coolant to the environment E.

Managing the heat of the internal combustion engine 110 is always a challenge. Accordingly, it may be advantageous to assist the heat exchanger 128 in dissipating heat generated by combustion occurring in the combustion chamber(s) of the internal combustion engine 110.

In the illustrated embodiment, the radiator 126 also includes at least one component 130 of the engine assembly 100. The component 130 is thermally disconnected from the heat exchanger 128. Herein, thermally disconnected means that the component 130 does not rely on a heat exchanger for cooling. In other words, in contrast to the internal combustion engine 110, the component 130 is not liquid cooled. Also, in other words, when the internal combustion engine 110 is turned on and combustion occurs in the chamber(s), the temperature of the engine 110, and more specifically the temperature of the liquid coolant exiting the engine 110, is higher than the temperature of the component 130. Therefore, heat transfer may occur from the liquid coolant that has been heated by the internal combustion engine 110 to the component 130. At least one of the primary components 130 has a primary function other than heat exchange. That is, the primary purpose of including at least one component 130 in the engine assembly is to assume a function unrelated to cooling/heating.

The component 130 may then dissipate heat received from the internal combustion engine 110 via the liquid coolant into the environment E. In a particular embodiment, the component 130 is air cooled by ambient air circulating within a cabin of an aircraft containing the engine assembly 100. In a particular embodiment, the cabin is an APU cabin of an aircraft.

In the illustrated embodiment, at least one component 130 includes two components, namely a gearbox 112 and an engine inlet 111. at least one component 130 may be any of the gearbox 112, the engine inlet 111, a compressor, a turbine, etc. in the illustrated embodiment, the component 130 is operatively connected to the internal combustion engine 110. it should be understood that the component 130 need not be operatively connected to the internal combustion engine 110. the component 130 may be any other component of the aircraft. for example, the component 130 may be an engine starter, a generator, a load compressor, an air filter, an actuator, a valve, any component having at least one movable portion, a line replacement unit (L RU), etc.

In the illustrated embodiment, the coolant circuit 116 includes a conduit 116a in heat exchange relationship with at least one component 130. The conduit 116a may be in contact with the component 130 such that the conduit 116a is thermally conductively coupled to the component 130. Herein, conductively connected means that heat is transferred between the conduit 116a and the component 130 via conduction. The conduit 116a may be wrapped around the member 130. The conduit 116a may simply be in contact with the component 130 or secured to the component 130 in any suitable manner, such as by welding. Alternatively or in combination, a channel for fluidly receiving a liquid coolant may be defined in the component 130. In other words, the cooling chambers and channels may be located around the component 130.

In certain embodiments, the conduit 116a may be selectively connected to the remainder of the coolant loop 116. In other words, the valve 132 may be located on the coolant loop 116. The valve 132 may operate in a first mode in which the valve 132 allows liquid coolant to circulate within the conduit 116b and a second mode in which the valve 132 prevents liquid coolant from flowing within the conduit 116 b. This may allow for selective use of the component 130 to assist the heat exchanger 128 in dissipating heat generated by the internal combustion engine 110 when desired. Having the ability to fluidly disconnect the conduit 116b from the remainder of the coolant circuit 116 may allow the pump to consume less energy because it does not have to overcome the pressure drop that may occur by circulating liquid coolant within the conduit 116 b.

In some flight phases, the internal combustion engine 110 must be exposed to very cold ambient temperatures for a long period of time before starting. It may be advantageous to warm up certain components of the engine assembly 110 prior to starting the internal combustion engine 110. More specifically, if the engine assembly 100 is an APU, the internal combustion engine 110 is shut down for substantially the entire cruise phase. When moving to the approach phase before landing, the power of the aircraft main engines is reduced and therefore they may not be able to provide all the electricity and compressed air required by the aircraft. The APU then starts to generate electricity and compressed air that were previously generated by the main engine during cruise. However, starting the engine 110 in this cold operating condition may require a relatively long warm-up time during which the efficiency of the engine assembly 110 is less than its nominal or steady-state efficiency. This may mean that the internal combustion engine 110 consumes more fuel during the warm-up phase than during the steady-state operating phase. It may be advantageous to reduce the duration of the warm-up phase to reduce the fuel consumption of engine 110.

In the illustrated embodiment, the engine assembly 100 includes a heat source 134 in heat exchange relationship with the coolant circuit 116 for heating the liquid coolant circulating therein, and as will be discussed below, for heating the component 130 prior to starting the engine 110. The heat source 134 may be an electric heater operatively connected to the power source S. The heat source 134 may be another heat exchanger of the aircraft. Any suitable heat source may be used without departing from the scope of the present disclosure.

During some other flight phases, the power of the internal combustion engine 110 must be shut down. However, when the internal combustion engine 110 has been operating for a long time, it may be very hot. After shutting down power, it may be advantageous to continue circulating liquid coolant to cool the internal combustion engine 110.

In the illustrated embodiment, the pump 118 may continue to operate using the electric motor 120 after the internal combustion engine 110 is powered off. Continuing to circulate liquid coolant after engine 110 is powered off may reduce cooling time as compared to configurations in which liquid coolant stops circulating in the coolant loop after engine 110 is powered off. Dissipating heat via both the component 130 and the heat exchanger 128 may reduce cooling time as compared to a configuration that uses only the heat exchanger 128 to dissipate heat of the internal combustion engine 110. In certain embodiments, reducing the cooling time extends the useful life of the engine. Reducing the cooling time may allow for reduced fuel consumption, since the engine does not have to run as long to achieve the same temperature reduction.

To operate the thermal management system 114, a liquid coolant is circulated. Heat is transferred from the internal combustion engine 110 to the liquid coolant. Heat is transferred from the liquid coolant to the environment E outside of the engine assembly 100 via both the heat exchanger 128 and the component 130 that is thermally disconnected from the heat exchanger 128.

In the illustrated embodiment, transferring heat to the environment E via the components 130 includes transferring heat to the gearbox 112 of the engine assembly 100. In the illustrated embodiment, transferring heat to the environment E via the component 130 includes transferring heat to the component 130 operably connected to the internal combustion engine 110.

In the depicted embodiment, circulating the liquid coolant includes actuating the pump 118 with one of the electric motor 120 and the engine shaft 110a of the internal combustion engine 110. Actuating the pump with the electric motor 118 may include disengaging the engine shaft 110a from the pump 118.

Referring now to fig. 2-3, a control system for controlling thermal management system 116 is shown generally at 200. Control system 200 includes a controller 210, controller 210 including a processor 212 and a computer-readable medium 214 operatively connected to processor 212 and having instructions stored thereon that are executable by processor 212 for controlling thermal management system 116. As shown, the controller 210 is operatively connected to an aircraft bus 211, and the aircraft bus 211 may be a 28V bus.

The controller 200 is configured for determining that the internal combustion engine 110 of the engine assembly 100 is off, circulating the liquid coolant in the coolant loop 116 for heating the liquid coolant, and circulating the heated liquid coolant toward the at least one component 130 thermally disconnected from the heat exchanger 128 of the engine assembly 100 for transferring heat from the heated liquid coolant to the at least one component 130.

In a particular embodiment, determining that the internal combustion engine 110 is off includes determining that the internal combustion engine 110 is cold. In this case, heating the liquid coolant includes heating the liquid coolant with the heat source 134. The heated liquid coolant may then be used to heat engine 110 and components 130.

In certain embodiments, determining that the internal combustion engine 110 is off includes determining that the internal combustion engine 110 is hot, or above a given temperature. In this case, heating the liquid coolant includes transferring heat from the internal combustion engine 110 to the liquid coolant.

In certain embodiments, determining that the internal combustion engine 110 is off further includes determining that the engine assembly 100 is below a given temperature. Alternatively, determining that the internal combustion engine 110 is off further includes determining that the flight phase of the aircraft is an approaching phase and that the engine 110 needs to be started.

In a particular embodiment, the controller 210 is operatively coupled to at least one sensor 216. As shown in fig. 2, the at least one sensor 216 includes a coolant temperature sensor 216a and a coolant pressure sensor 216b, both of which may be operatively connected to an Engine Control Unit (ECU) 218. As shown, the engine control unit 218 is operatively connected to the controller 210. A coolant temperature sensor 216a may be used to monitor the temperature of the component 130 and a coolant pressure sensor 216b may be used to monitor the pressure of the coolant. If one or both of the monitored temperature and the monitored pressure reach respective given thresholds, the controller 210 may initiate circulation of the liquid coolant and actuation of the heat source 134 to preheat the component 130.

As shown on fig. 2, the engine control unit 218 is operatively connected to a power source S. As shown, the controller 210 is operatively connected to a power source S via an engine control unit 218.

The valve 132 is operatively connected to the engine control unit 218 and the controller 210. In other words, valve 132 is operatively connected to controller 210 via engine control unit 218. In the illustrated embodiment, the pump 118 is electrically driven to turn pumping on/off whenever the engine 110 is shut down to warm up or cool down hardware (e.g., the engine 110). The valve or control valve 132 may provide the possibility to allow coolant fluid to flow in a particular engine module as needed before engine start-up, in operation, and after shutdown. Depending on the temperature and pressure of the coolant and the engine state, the ECU 218 may have logic to open the required cooling passages. Multiple temperature sensing points may be used.

In one embodiment, the control system 200 may monitor critical component temperatures during flight and turn on the pump 118 for preheating as needed.

In a particular embodiment, depending on the application, the cooling pump 118 may be driven by a direct mechanical drive shaft that takes over the electric motor 120 when the engine 110 is started using the overrunning clutch 124 to engage and disengage the drive shaft.

In certain embodiments, having the ability to warm up or cool down components such as the gearbox 112, the engine inlet 111, and the engine 110 itself may allow for reduced cooling time after shutting down the engine 110 when the engine 110 is powered off, which may allow for fuel savings. It is possible to allow the time for warming up the engine 110 before starting it to be reduced. When starting the engine 110, thermal stress and wear of mechanical components of the engine (e.g., gears, bearings, etc.) may be reduced. Drag may be reduced and lubrication of engine components may be improved during warm-up. The system 114 may be a single system for warming and/or cooling different engine systems, such as a transmission, a casing of the engine 110, and the like.

The above description is intended to be exemplary only, and those skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, the thermal management systems disclosed herein may be used in turboshafts, turboprops, and turbofan engines. Still other modifications that fall within the scope of the invention will be readily apparent to those skilled in the art in view of a review of this disclosure, and such modifications are intended to fall within the scope of the appended claims.