Engine assembly with a thermal management system

Abstract

An engine assembly for an aircraft is disclosed, which includes: a combustion engine including a coolant circuit in heat exchange relationship with a radiator, the radiator including a heat exchanger and at least one component having a primary function different from heat exchange. A method of operating a system is also disclosed.

Description

Technical Field

[0001] The present application generally relates to aircraft engines and, more particularly, to systems and methods for managing heat generated by such engines. Background Art

[0002] The performance of an aircraft engine during startup is not as good as during steady-state operation because various engine components and fluids are not at their optimal operating temperatures. After the aircraft engine is shut down, a significant amount of time elapses before the engine cools. Additionally, providing sufficient cooling to the hot components of an aircraft engine is a long-term challenge. Therefore, improvements are appropriate. Summary of the Invention

[0003] In one aspect, there is provided an engine assembly for an aircraft, comprising: a combustion engine including a coolant circuit in thermal exchange relationship with a radiator, the radiator including a heat exchanger and at least one component having a primary function different from heat exchange.

[0004] In another aspect, there is provided a method of operating a thermal management system of an engine assembly for an aircraft, 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 a heat exchanger and a component having a primary function different from heat exchange.

[0005] In yet another aspect, there is provided a method of operating a thermal management system of an engine assembly for an aircraft, the engine assembly including a combustion engine liquid-cooled with a liquid coolant, the method comprising: determining that the combustion engine of the engine assembly is shut down; circulating the liquid coolant in the coolant circuit for heating the liquid coolant; and circulating the heated liquid coolant toward at least one component having a primary function different from heat exchange for transferring heat from the heated liquid coolant to the at least one component. Brief Description of the Drawings

[0006] Reference is now made to the drawings, in which:

[0007] Figure 1 is a schematic cross-sectional view of a rotary internal combustion engine according to a particular embodiment;

[0008] Figure 2 is a schematic view of an engine assembly according to one embodiment; and

[0009] Figure 3 is for Figure 2 a schematic view of a control system of an engine assembly. Detailed Description

[0010] Reference Figure 1 shows schematically a rotary internal combustion engine 10 known as a Wankel engine. 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 chamber 20. The inner surface of the peripheral wall 18 of the chamber 20 has a profile defining two convex corners, which is preferably an epitrochoid.

[0011] An inner body or rotor 24 is received within the chamber 20. The rotor 24 has axially spaced end faces 26 adjacent to the outer body end walls 14, and a peripheral face 28 extending therebetween. The peripheral face 28 defines three circumferentially spaced vertex portions 30, and has a generally triangular profile with outwardly arched sides 36. The vertex portions 30 are in sealing engagement with the inner surface of the peripheral wall 18 to form three rotary combustion chambers 32 between the inner rotor 24 and the outer body 12. The geometric axis of the rotor 24 is offset and parallel to the axis of the outer body 12.

[0012] The combustion chambers 32 are sealed. In the illustrated embodiment, each rotor vertex portion 30 has a vertex seal 52 that extends from one end face 26 to the other and is radially outwardly biased against the peripheral wall 18. End seals 54 engage each end of each vertex seal 52 and are biased against the respective end walls 14. Each end face 26 of the rotor 24 has at least one arcuate face seal 60 that extends from each vertex portion 30 to each adjacent vertex portion 30, adjacent to the rotor periphery throughout its length but inside thereof, in sealing engagement with the end seals 54 adjacent to each of its ends, and is biased into sealing engagement with the adjacent end walls 14. Alternative sealing arrangements are also possible.

[0013] Although not shown in the drawings, the rotor 24 is journal supported on an eccentric portion of a shaft such that rotation of the shaft rotates the rotor 24 to effect an orbital revolution within the stator chamber 20. As the rotor 24 moves around the stator chamber 20, for each complete rotation of the rotor 24, the shaft rotates three times. Oil seals are provided around the eccentric to prevent lubricating oil from leaking radially outward between the respective rotor end faces 26 and the outer body end walls 14. During each rotation of the rotor 24, the volume of each chamber 32 changes and moves around the stator chamber 20 to undergo four stages of intake, compression, expansion, and exhaust, which are analogous to the strokes in a reciprocating internal combustion engine having a four-stroke cycle.

[0014] The engine includes a main 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. Ports 40, 42, 44 may be defined in an end wall 14 of a peripheral wall 18. In the illustrated embodiment, the inlet port 40 and the purge port 42 are defined in the end wall 14 and communicate with the same intake duct 34, which is defined as a passage in the end wall 14, and the exhaust port 44 is defined through the peripheral wall 18. Alternative configurations are possible.

[0015] In a particular embodiment, 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 fuel-rich air mixture near an ignition source and a leaner mixture at other locations, 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, the rotary engine 10 operates on the principle of the Miller or Atkinson cycle through appropriate relative positions of the main inlet port 40 and the exhaust port 44, having a compression ratio lower than its expansion ratio.

[0016] Now referring Figure 2 , an engine assembly is generally shown 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 may be the rotary engine 10 described above with reference Figure 1 to herein. 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 a particular embodiment, 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. Patent No. 7,753,036 to Lents et al. published on July 13, 2010, or such as described in U.S. Patent No. 7,775,044 to Julien et al. published on August 17, 2010, or such as described in U.S. Patent Publication No. 2015 / 0275749 to Thomassin et al. published on October 1, 2015, or such as described in U.S. Patent Publication No. 2015 / 0275756 to Bolduc et al. published on October 1, 2015, the entire contents of which are incorporated herein by reference.

[0017] The internal combustion engine 110 is fluidly connected to an engine inlet 111, which is fluidly connected to the environment E external to the engine assembly 100.

[0018] In the illustrated embodiment, engine assembly 100 includes a gearbox 112 drivingly engaged with an internal combustion engine 110. More specifically, internal combustion engine 110 has an engine shaft 110a connected to an input 112a of gearbox 112. Gearbox 112 may be connected to a plurality of components to transfer rotational input from engine shaft 110a. Gearbox 112 may produce a rotational speed ratio between its input 112a and its output 112b.

[0019] In the illustrated embodiment, engine assembly 100 includes a thermal management system 114. Thermal management system 114 includes a coolant circuit 116 configured to circulate a liquid coolant. Coolant circuit 116 is in a heat exchange relationship with internal combustion engine 110. In a particular embodiment, internal combustion engine 110 includes a housing, such as Figure 1 the peripheral wall 18 of a rotary engine, which defines a conduit therein; the conduit is fluidly connected to coolant circuit 114. The liquid coolant may be able to cool internal combustion engine 110 by absorbing heat from the housing of internal combustion engine 110 via convection. Accordingly, the temperature of the liquid coolant increases after it passes through the housing.

[0020] To cause flow of the liquid coolant within coolant circuit 116, engine assembly 100 includes a pump 118 fluidly connected to coolant circuit 116. Pump 118 may be drivingly engaged by internal combustion engine 110 and / or by an electric motor 120 powered by a power source S. In a particular embodiment, power source S is a battery 122. Alternatively, power source S may be a generator drivingly engaged by internal combustion engine 110 or another engine of an aircraft incorporating engine assembly 100.

[0021] In the illustrated embodiment, thermal management system 114 includes an expansion tank 117 fluidly connected to coolant circuit 116. Expansion tank 117 may be pressurized to a given pressure (which may be 35 PSI) and may be used for disposal of thermally induced volume changes of the liquid coolant.

[0022] In some cases, it may be useful to circulate the liquid coolant in coolant circuit 116 when internal combustion engine 110 shuts down power, stalls, or shuts off. Herein, shutting down power, stalling, or shutting off means that internal combustion engine 110 does not cause rotation of engine shaft 110a. In other words, by shutting down power, stalling, or shutting off, combustion does not occur in the (one or more) combustion chambers of internal combustion engine 110.

[0023] In the illustrated embodiment, pump 118 is selectively drivingly engaged by internal combustion engine 110 or electric motor 120. In the depicted embodiment, pump 118 is drivingly engaged by internal combustion engine via gearbox 112.

[0024] To allow the pump 118 to engage with the electric motor 120, the engine assembly 100 includes a clutch 124 having an input portion 124a and an output portion 124b. The input portion 124a is drivingly engaged with the internal combustion engine 110 via a gearbox 112, and the output portion 124b is drivingly engaged with the pump 118. The clutch 124 is operable in a first configuration in which the input portion 124a is drivingly engaged with the output portion 124b, and the clutch 124 is operable in a second configuration in which the input portion 124a is disengaged from the output portion 124b. 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 disengaged from the pump 118. In the second configuration, the clutch 124 allows the electric motor 120 to drive the pump 118 without having to overcome the load generated by the shutdown of the internal combustion engine 110.

[0025] In a particular embodiment, the clutch 124 is a sprag clutch. In a sprag clutch, if the input portion 124a rotates at a speed greater than the output portion 124b, the input portion 124a is drivingly engaged with the output portion 124b, and if the output portion 124b rotates at a speed greater than the input portion 124a, the input portion 124a allows the output portion 124b to rotate independently of the input portion 124a.

[0026] The thermal management system 114 further includes a radiator 126 in a heat exchange relationship with the coolant circuit 116. In the illustrated embodiment, the radiator 126 includes a heat exchanger 128. Heat generated by the engine 110 must be dissipated in the environment E. The heat exchanger 128 can be used for this purpose.

[0027] 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 a heat exchange relationship with the at least one first conduit 128b. The at least one first conduit 128a is fluidly connected to the coolant circuit 116, and the at least one second conduit 128b is fluidly connected to the environment E. A blower B can be used to direct air from the environment E through the at least one second conduit 128b of the heat exchanger 128. In a particular embodiment, the blower B can be replaced by a scoop located on the outer surface of the aircraft. The temperature of the air in the environment E is generally 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.

[0028] Managing the heat of the internal combustion engine 110 is always a challenge. Thus, it may be advantageous to assist the heat exchanger 128 in dissipating the heat generated by combustion occurring in the (multiple) combustion chambers of the internal combustion engine 110.

[0029] 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 the heat exchanger for cooling. In other words, contrary 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 (multiple) chambers, the temperature of the engine 110, more specifically the temperature of the liquid coolant leaving the engine 110, is higher than the temperature of the component 130. Thus, heat transfer may occur from the liquid coolant that has been heated by the internal combustion engine 110 to the component 130. The at least one main component 130 has a main function different from heat exchange. That is, the main purpose of including at least one component 130 in the engine assembly is to perform a function unrelated to cooling / heating.

[0030] Then, the component 130 can dissipate the heat received from the internal combustion engine 110 via the liquid coolant to the environment E. In a particular embodiment, the component 130 is air-cooled by ambient air circulated within a compartment of an aircraft that houses the engine assembly 100. In a particular embodiment, the compartment is the APU compartment of the aircraft.

[0031] In the illustrated embodiment, the at least one component 130 includes two components, namely the transmission 112 and the engine inlet 111. The at least one component 130 can be any one of the transmission 112, the engine inlet 111, a compressor, a turbine, etc. In the illustrated embodiment, the component 130 is operably connected to the internal combustion engine 110. It should be understood that the component 130 does not need to be operably connected to the internal combustion engine 110. The component 130 can be any other component of the aircraft. For example, the component 130 can be an engine starter, a generator, a load compressor, an air filter, an actuator, a valve, any component having at least one movable part, a line replaceable unit (LRU), etc. In a particular embodiment, having at least one component 130 as the engine inlet 111 allows de-icing or prevents ice accumulation on the inlet 111. All of the above-listed possibilities for the at least one component 130 have a main function different from heat exchange. For example, the main function of an engine starter is to start the engine, the main function of a generator is to generate electricity, the main function of a load compressor is to compress air for the cabin of an aircraft equipped with the engine assembly, the main function of an actuator is to move another component, the main function of a valve is to allow or prevent fluid communication between two components, the main function of a compressor is to compress air before feeding it to a combustion engine, and the main function of a turbine is to extract energy from the combustion gases leaving the combustion engine.

[0032] In the illustrated embodiment, the coolant circuit 116 includes a conduit 116a that is in a heat exchange relationship with at least one component 130. The conduit 116a can be in contact with the component 130 such that the conduit 116a is thermally connected to the component 130. As used herein, being thermally connected means that heat is transferred between the conduit 116a and the component 130 via conduction. The conduit 116a can be wrapped around the component 130. The conduit 116a can simply be in contact with the component 130 or fixed to the component 130 in any suitable manner, such as by welding. Alternatively or in combination, channels can be defined in the component 130 for fluidly receiving a liquid coolant. In other words, the cooling chamber and channels can be located around the component 130.

[0033] In a particular embodiment, the conduit 116a can be selectively connected to the remainder of the coolant circuit 116. In other words, a valve 132 can be located on the coolant circuit 116. The valve 132 can operate in a first mode and a second mode. In the first mode, the valve 132 allows liquid coolant to circulate within the conduit 116b. In the second mode, the valve 132 prevents liquid coolant from flowing within the conduit 116b. This can allow the component 130 to be selectively used, when needed, to assist the heat exchanger 128 in dissipating heat generated by the internal combustion engine 110. Having the ability to fluidly disconnect the conduit 116b from the remainder of the coolant circuit 116 can 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 116b.

[0034] During some flight phases, the internal combustion engine 110 must be started after being exposed to very cold ambient temperatures for an extended period of time. It may be advantageous to preheat certain components of the engine assembly 110 before starting the internal combustion engine 110. More specifically, if the engine assembly 100 is an APU, the internal combustion engine 110 is shut down during substantially the entire cruise phase. When transitioning to the approach phase before landing, the power of the aircraft main engines is reduced and thus they may not be able to provide all of the electrical power and compressed air required by the aircraft. Then, the APU is started to generate electrical power and compressed air that were previously generated by the main engines during cruise. However, starting the engine 110 under such cold operating conditions may require a relatively long preheating 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 preheating phase compared to during the steady-state operation phase. It may be advantageous to reduce the duration of the preheating phase to reduce the fuel consumption of the engine 110.

[0035] In the illustrated embodiment, the engine assembly 100 includes a heat source 134 in thermal exchange relationship with a coolant circuit 116 for heating a liquid coolant circulating therein and, as will be discussed below, for heating components 130 prior to starting engine 110. The heat source 134 can be an electric heater operably connected to a power source S. The heat source 134 can be another heat exchanger of the aircraft. Any suitable heat source can be used without departing from the scope of the present disclosure.

[0036] 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 the power, it may be advantageous to continue circulating the liquid coolant to cool the internal combustion engine 110.

[0037] In the illustrated embodiment, after the internal combustion engine 110 shuts down power, the pump 118 can continue to operate using the electric motor 120. Compared with a configuration where the liquid coolant stops circulating in the coolant circuit after the engine 110 shuts down, continuing to circulate the liquid coolant after the engine 110 shuts down power may reduce the cooling time. Compared with a configuration that only uses the heat exchanger 128 to dissipate the heat of the internal combustion engine 110, dissipating heat via both the component 130 and the heat exchanger 128 can reduce the cooling time. In a particular embodiment, reducing the cooling time extends the service life of the engine. Reducing the cooling time may allow for a reduction in fuel consumption because the engine does not have to run as long to achieve the same temperature reduction.

[0038] To operate the thermal management system 114, the 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 the engine assembly 100 via both the heat exchanger 128 and the component 130 that is thermally disconnected from the heat exchanger 128.

[0039] In the illustrated embodiment, transferring heat to the environment E via the component 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.

[0040] In the depicted embodiment, circulating the liquid coolant includes actuating the pump 118 with either the electric motor 120 or the engine shaft 110a of the internal combustion engine 110. Actuating the pump with the electric motor 118 can include disengaging the engine shaft 110a from the pump 118.

[0041] Now refer to Figure 2-3, generally shown at 200 is a control system for controlling a thermal management system 116. The control system 200 includes a controller 210, which includes a processor 212 and a computer-readable medium 214 operably connected to the processor 212 and having instructions stored thereon that can be executed by the processor 212 to control the thermal management system 116. As shown, the controller 210 is operably connected to an aircraft bus 211, which can be a 28V bus.

[0042] The controller 200 is configured to determine that the internal combustion engine 110 of the engine assembly 100 is off, circulate a liquid coolant in the coolant loop 116 to heat the liquid coolant, and circulate the heated liquid coolant toward at least one component 130 that is thermally disconnected from the heat exchanger 128 of the engine assembly 100 to transfer heat from the heated liquid coolant to the at least one component 130.

[0043] In a particular embodiment, determining that the internal combustion engine 110 is off includes determining that the internal combustion engine 110 is cold. In such a case, heating the liquid coolant includes heating the liquid coolant with a heat source 134. Then, the heated liquid coolant can be used to heat the engine 110 and the component 130.

[0044] In a particular embodiment, 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 such a case, heating the liquid coolant includes transferring heat from the internal combustion engine 110 to the liquid coolant.

[0045] In a particular embodiment, 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 the approach phase and the engine 110 needs to be started.

[0046] In a particular embodiment, the controller 210 is operably connected to at least one sensor 216. As Figure 2 shown, the at least one sensor 216 includes a coolant temperature sensor 216a and a coolant pressure sensor 216b, both of which can be operably connected to an engine control unit (ECU) 218. As shown, the engine control unit 218 is operably connected to the controller 210. The coolant temperature sensor 216a can be used to monitor the temperature of the component 130, while the coolant pressure sensor 216b can be used to monitor the pressure of the coolant. If one or both of the monitored temperature and the monitored pressure reach corresponding given thresholds, the controller 210 can start the circulation of the liquid coolant and the actuation of the heat source 134 to preheat the component 130.

[0047] AsFigure 2 As shown above, the engine control unit 218 is operably connected to the power source S. As shown, the controller 210 is operably connected to the power source S via the engine control unit 218.

[0048] The valve 132 is operably connected to the engine control unit 218 and the controller 210. In other words, the valve 132 is operably connected to the controller 210 via the engine control unit 218. In the illustrated embodiment, the pump 118 is electrically driven at any time when the engine 110 is shut down to turn on / off pumping, thereby preheating or cooling the hardware (e.g., the engine 110). The valve or control valve 132 can provide the possibility to allow the coolant fluid to flow in specific engine modules as needed before engine startup, during operation, and after shutdown. Depending on the temperature and pressure of the coolant and the engine state, the ECU 218 may have the logic to open the required cooling channels. Multiple temperature sensing points can be used.

[0049] In one embodiment, the control system 200 can monitor the critical component temperatures during flight and turn on the pump 118 for preheating as needed.

[0050] In a particular embodiment, depending on the application, the coolant pump 118 can be driven by a direct mechanical drive shaft that takes over from the electric motor 120 when starting the engine 110 using the overrunning clutch 124 to engage and disengage the drive shaft.

[0051] In a particular embodiment, when the engine 110 shuts off power, the ability to preheat or cool components such as the transmission 112, the engine inlet 111, and the engine 110 itself can allow for a reduction in the cooling time after shutting down the engine 110, which can allow for fuel savings. It can allow for a reduction in the time to preheat the engine 110 before starting it. When starting the engine 110, the thermal stress and wear on the mechanical components of the engine (e.g., gears, bearings, etc.) can be reduced. The drag can be reduced and the lubrication of the engine components can be improved during preheating. The system 114 can be a single system for preheating and / or cooling different engine systems such as the transmission, the housing of the engine 110, etc.

[0052] The above description is only intended to be exemplary, and those skilled in the art will recognize that changes can be made to the described embodiments without departing from the scope of the disclosed invention. For example, the thermal management system disclosed herein can be used in turboprop, turbopropeller, and turbofan engines. In view of the review of this disclosure, still other modifications that fall within the scope of the present invention will be apparent to those skilled in the art, and these modifications are intended to fall within the scope of the appended claims.

Claims

1. An engine assembly for an aircraft, which comprises: A combustion engine including a coolant circuit in thermal exchange relationship with a radiator, the radiator including a heat exchanger and at least one other component of the engine assembly, the at least one other component having a primary function different from heat exchange. Wherein, the coolant circuit includes a conduit and a valve is located on the coolant circuit, the valve having a first mode and a second mode. In the first mode, the valve allows liquid coolant to circulate within the conduit; in the second mode, the valve prevents the liquid coolant from flowing within the conduit. The valve is configured to selectively use the at least one other component to assist the heat exchanger in dissipating heat generated by the combustion engine when needed. Wherein, the coolant circuit includes a pump, the pump is selectively driven and engaged by an electric motor operably connected to a power source or by the engine shaft of the combustion engine. After the combustion engine shuts off power, the pump continues to operate using the electric motor. Heat is transferred from the combustion engine to the liquid coolant and is transferred from the liquid coolant to the environment outside the engine assembly via both the heat exchanger and at least one other component that is thermally disconnected from the heat exchanger. Wherein, the at least one other component includes two components, a gearbox and an engine inlet, and the engine inlet allows de-icing or prevents ice accumulation on the engine inlet.

2. The engine assembly according to claim 1, wherein, The coolant circuit includes a conduit in contact with the at least one other component, and the conduit is thermally connected to the at least one other component.

3. The engine assembly according to claim 1, wherein, The coolant circuit includes a conduit in thermal exchange relationship with the at least one other component, and the conduit is selectively connectable to the remainder of the coolant circuit.

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

5. The engine assembly according to claim 1, wherein, The at least one other component is operably connected to the combustion engine.

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

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

8. The engine assembly according to claim 1, further comprising a clutch having an input portion drivingly engaged by the engine shaft and an output portion drivingly engaged with the pump. The clutch is operable in a first configuration and a second configuration. In the first configuration, the input portion is drivingly engaged with the output portion; in the second configuration, the input portion is disengaged from the output portion.

9. The engine assembly according to claim 1, wherein the combustion engine is a reciprocating engine.

10. A method of operating a thermal management system of an engine assembly according to any one of claims 1-9 for an aircraft, the method comprising: circulating a liquid coolant; transferring heat from the 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 a heat exchanger and a component whose primary function is different from heat exchange.

11. The method according to claim 10, wherein transferring heat to the environment via the component includes transferring heat to a gearbox of the engine assembly.

12. The method according to claim 10, wherein transferring heat to the environment via the component includes transferring heat to the component operably connected to the combustion engine.

13. The method according to claim 10, wherein circulating the liquid coolant includes actuating a pump with one of an electric motor and an engine shaft of the combustion engine.

14. The method according to claim 10, actuating the pump with the electric motor includes disengaging the engine shaft from the pump.

15. A method of operating a thermal management system of an engine assembly according to any one of claims 1-9 for an aircraft, the engine assembly including a combustion engine liquid-cooled by a liquid coolant, the method comprising: determining that the combustion engine of the engine assembly is shut down; circulating the liquid coolant in a coolant circuit for heating the liquid coolant; and circulating the heated liquid coolant towards 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.

16. The method according to claim 15, wherein determining that the combustion engine is shut down includes determining that the combustion engine is cold, and heating the liquid coolant includes heating the liquid coolant with a heat source.

17. The method according to claim 15, wherein determining that the combustion engine is shut down includes determining that the combustion engine is hot, and heating the liquid coolant includes transferring heat from the combustion engine to the liquid coolant.

18. The method according to claim 15, wherein determining that the combustion engine of the engine assembly is shut down further includes determining that the engine assembly is below a given temperature.

19. The method according to claim 15, wherein determining that the combustion engine of the engine assembly is shut down further includes determining that the flight phase of the aircraft is an approach phase.

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