GB2594072A

GB2594072A - Hybrid aircraft propulsion system

Info

Publication number: GB2594072A
Application number: GB2005490.4A
Authority: GB
Inventors: P Scothern David
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-04-15
Filing date: 2020-04-15
Publication date: 2021-10-20
Also published as: GB202005490D0

Abstract

An aircraft propulsion system (fig.2,5) has an internal combustion engine (fig.2,10) such as a gas turbine engine; an electric motor (fig.2,28); and a propulsor (fig.2,12) driven by at least the motor. The motor has an engine fuel cooled cooling system 40 having a heat exchanger 46 which transfers heat from the motor to engines fuel. The system also has an overboard fuel dump system with a drain valve 54 provided downstream of the heat exchanger, preferably also a number of fuel injectors 17 downstream of the valve. The system may be a series hybrid propulsion system, in which the engine is coupled to an electric generator 32 providing power to the motor, and the engine is not mechanically coupled to a propulsor. Alternatively, the system may be a parallel hybrid propulsion system where the engine is coupled to a propulsor. An electrical energy storage device (fig.2,30) for the motor may be provided. Preferably, the engine has an oil cooling with a fuel-oil heat exchanger 50 downstream of the motor heat exchanger, rejecting heat from engine oil to engine fuel. The motor cooling system preferably has an intermediate cooling loop with a second heat exchanger 48 using an intermediate cooling fluid.

Description

HYBRID AIRCRAFT PROPULSION SYSTEM

The present disclosure concerns an aircraft comprising a hybrid propulsion system for an aircraft and an aircraft comprising the hybrid propulsion system.

Hybrid aircraft have been proposed, in which an internal combustion engine is combined with one or more electric motors to drive one or more propulsors. Parallel hybrid systems can be distinguished from so-called "series hybrid" systems, in that in a parallel hybrid system, a mechanical connection is provided between the internal combustion engine and at least one propulsor, with at least one electric motor driving either the same propulsor as that driven by the internal combustion engine, or a further propulsor. In series hybrid systems, the internal combustion engine is not mechanically coupled to a propulsor.

According to a first aspect there is provided an aircraft propulsion system comprising; an internal combustion engine; an electric propulsor motor; a propulsor configured to be driven by at least the propulsor motor; wherein the electric motor comprises an internal combustion engine fuel cooled cooling system comprising a motor heat exchanger configured to transfer heat from the electric motor to internal combustion engine fuel, and wherein the propulsion system further comprises an overboard fuel dump system comprising an overboard drain valve provided downstream in fuel flow of the motor heat exchanger.

Advantageously, in the event of a failure which requires an in-flight diversion, fuel can be dumped from the aircraft, allowing for a lower landing weight, which increases safety. The provision of the overboard drain valve downstream of the heat exchanger permits continued operation of the heat exchanger during fuel dumping operations, thereby maintaining cooling levels of the electric motor. In addition, in the event of an emergency requiring additional motor power, higher cooling levels can be accommodated, by operation of the overboard drain valve.

The internal combustion engine may comprise one or more fuel injectors downstream of the overboard drain valve.

The aircraft propulsion system may comprise a series hybrid propulsion system, in which the internal combustion engine is coupled to an electric generator configured to provide electrical power to the propulsor motor, and wherein the internal combustion engine is not mechanically coupled to a propulsor.

Alternatively, the aircraft propulsion system may comprise a parallel hybrid propulsion system, in which the internal combustion engine is mechanically coupled to at least one propulsor. Where the aircraft propulsion system comprises a parallel hybrid system, the internal combustion engine may be coupled to an electric generator configured to provide electrical power to the propulsor motor.

The aircraft propulsion system may comprise an energy storage device configured to store electrical energy for provision to the propulsor motor.

The internal combustion engine may comprise a gas turbine engine.

The internal combustion engine may comprise a cooling system configured to utilise internal combustion engine fuel as a heat sink.

The internal combustion engine cooling system may comprise an oil cooling system, which may comprise a fuel-oil heat exchanger configured to reject heat from the internal combustion engine oil to the internal combustion engine fuel.

The fuel-oil heat exchanger may be provided downstream in fuel flow of the motor heat exchanger.

The motor cooling system may comprise an intermediate cooling loop, comprising a second heat exchanger configured to exchange heat between the electric motor and an intermediate cooling fluid, the motor heat exchanger being configured to exchange heat between the intermediate cooling fluid and the fuel.

The internal combustion engine fuel may comprise a liquid fuel such as a liquid hydrocarbon. Alternatively, the internal combustion engine may comprise a gaseous fuel, such as a gaseous hydrocarbon or gaseous hydrogen.

According to a second aspect there is provided an aircraft comprising a propulsion system according to the first aspect.

The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein.

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a plan view of a first aircraft comprising a parallel hybrid propulsion system; Figure 2 is a schematic diagram of a parallel hybrid propulsion system for the aircraft of figure 1; Figure 3 is a schematic diagram of a fuel and cooling system of the propulsion system of figure 2; Figure 4 is a plan view of a second aircraft comprising a series hybrid propulsion system; and Figure 5 is a schematic diagram of a fuel and cooling system of the propulsion system of figure 4.

With reference to Figure 1, an aircraft 1 is shown. The aircraft is of conventional configuration, having a fuselage 2, wings 3, tail 4 and a pair of propulsion systems 5. One of the propulsion systems 5 is shown in figure detail in figure 2.

Figure 2 shows the propulsion system 5 schematically. The propulsion system 5 includes an internal combustion engine in the form of a gas turbine engine 10. The gas turbine engine 10 comprises, in axial flow series, a propulsor in the form of a fan 12, a compressor 14, combustion equipment 16 and high and low-pressure turbines 18, 20.

The gas turbine engine 10 works in the conventional manner so that air is accelerated by the fan 12 to produce two air flows: a first core air flow into the compressor 14 and a second air flow which bypasses the compressor 14 to provide propulsive thrust. The core air flows through the compressor 14 where it is compressed, before delivering that air to the combustion equipment 16, where it is mixed with fuel and the mixture combusted. Fuel is supplied to the combustors 16 via fuel injectors 17. The resultant hot combustion products then expand through, and thereby drive the turbines 18, 20 before being exhausted through a nozzle to provide additional propulsive thrust. The high 18 and low-pressure turbines 18, 20 drive respectively the compressor 14 and fan 12, each by suitable interconnecting shaft 22, 24.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. three) and/or an alternative number of compressors and/or turbines. Further, the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

The propulsion system 5 further comprises one or more electrical machines driving one or more propulsors. In particular, the system 5 comprises an electric motor 28. The motor 28 is of a conventional type, such as an induction or permanent magnet electric machine, and is configured to drive a propulsor such as the fan 12. In the present embodiment, the motor 28 is coupled to the fan 12 via the low-pressure shaft 24. In this embodiment, the electric motor 28 is of a "core shaft mounted" type, in which a rotor 29 of the motor 28 is mounted directly to a surface of the low-pressure shaft 24, and is surrounded by a stator 31, provided radially outwardly of the rotor 29. The stator comprises electrical windings (not shown), which can be energised to produce a rotating magnetic field. This rotating magnetic field interacts with a magnetic field of the rotor 29, to cause rotation when acting as a motor. Consequently, the fan 12 may be powered by either or both of the gas turbine engine 10 via the low-pressure turbine 20, and the motor 28.

The electric motor 28 is coupled to an energy storage device 30 in the form of one or more of a chemical battery, fuel cell, and capacitor, which provides the electric motor 28 with electrical power during operation. In some cases, multiple energy storages systems, which may be of different types (chemical battery, fuel cell etc) may be provided for each propulsion system 5. In other cases, a common energy storage device 30 may be provided for multiple propulsion systems.

The propulsion system optionally further comprises a generator 32, which is coupled to one or both of the motor 28 and the energy storage device 30, such that additional electrical energy can be provided in operation. The generator 32 is typically driven by the low-pressure shaft 24 of the gas turbine engine 10. The generator 32 may be coupled to the shaft 24 via a gearbox and / or clutch to allow for selectively connecting and disconnecting the generator 32 from the shaft 24. In some cases, the motor 28 may act as a generator.

A controller 34 is provided, which is configured to control at least the motor 28 and energy storage device 30, to control the torque provided by the motor 28, and the charging / discharging of the energy storage device 30. The controller 34 may also be configured to control operation of the generator 32, to control electrical power produced by the generator 32.

Figure 3 is a schematic, illustrating a fuel and cooling system 40 for the propulsion system of figure 2. The system comprises a fuel line 42, which is configured to transfer fuel from a fuel tank 44 to various systems. The fuel is a fluid, and is of any type suitable for use with an internal combustion engine. Typically, the fuel is a liquid hydrocarbon fuel such as mineral or synthetic jet fuel. Alternatively, the fuel could be a non-hydrocarbon fuel such as ammonia or liquid hydrogen. Similarly, the fuel could be in a gaseous form. Suitable fuels include gaseous hydrocarbons such as methane or propane, or non-hydrocarbon fuels such as gaseous hydrogen.

In use, fuel is provided from the fuel tank 44 to the fuel injectors 17 of the engine 10. In between the fuel tank 44 and the fuel injectors 17, are a number of cooling systems that utilise the fuel as a cooling fluid The first cooling system element is a motor heat exchanger 46. This is provided downstream of the fuel tank 44, and upstream of further components of the cooling system. In this embodiment, the motor heat exchanger 46 comprises a liquid-to-liquid heat exchanger, which transfer heat from an intermediate cooling fluid (such as oil) in a cooling line 58 to the fuel in the fuel line 42. Upstream of the motor heat exchanger 46 in intermediate cooling fluid flow is a second heat exchanger 48, which exchanges heat between components of the motor 28, and the intermediate cooling fluid. The intermediate cooling fluid flows around a closed loop, such that the intermediate fluid continuously flows between the motor and second heat exchangers 46, 48.

Typically, one or more components of the motor 28 may need active cooling via the heat exchanger 48. In particular, "active components" of the motor 28 such as components through which current flows, may require cooling. For instance, the stator 31 (and in particular, stator windings) may require active cooling. In one embodiment therefore, the second heat exchanger 48 comprises one or more of cooling tubes which are provided in close proximity to stator 31 components, such as electrical windings and laminations (not shown).

Further components of the motor 28 that may require cooling via the intermediate cooling fluid include the rotor 29, as well as power electronics devices such as the motor controller 30.

In other embodiments, the second heat exchanger may be omitted, and the fuel in the fuel line 42 may instead be used to directly cool motor components. In such a case, the motor heat exchanger 46 may comprise cooling lines, which are provided in close proximity to cooled motor components. Such a system will reduce the weight associated with the intermediate cooling fluid and second heat exchanger, but may not be suitable in some circumstances, due to the increased risk of fire.

The cooling system optionally further comprises an fuel-oil heat exchanger 50. The fuel-oil heat exchanger 50 comprises a heat exchanger configured to transfer heat from oil of the engine 10 oil system (not shown) to the fuel in the fuel line 42. The fuel-oil heat exchanger 50 is provided in the fuel line 42 between the fuel tank 44 and the fuel injectors 17 in fuel flow, and in this embodiment is provided between the motor heat exchanger 46 and the fuel injectors 17.

The system further comprises a fuel dump system. The fuel dump system is provided to allow for fuel to be dumped overboard the aircraft in an emergency, i.e. from the fuel tank 44 to the airstream, bypassing the engine 10 and fuel injectors 17.

The fuel dump system comprises a fuel dump nozzle 52, which forms part of the fuel line 42, and carries fuel to the external airstream in use. Upstream of the nozzle is a overboard fuel drain valve 54, which is configurable between open and closed positions. The fuel drain valve 54 is provided as part of a bypass line 56, which forms an offshoot of the fuel line 42, and provides a second flow path, which flows from the fuel tank to the airstream, bypassing the fuel injector 17. Consequently, when the valve 54 is closed, fuel flows from the fuel tank 44, through the motor heat exchanger 46 and oil-fuel heat exchanger 50. When the valve 54 is closed, fuel flow from the fuel line 42, to the dump nozzle 52, and out of the aircraft. A further valve (not shown) may be provided to prevent fuel from reaching the fuel injectors 17. In some cases, the valve 54 may be provided as part of the fuel line, and may be configured to switch flows between the fuel injectors 17 and the dump nozzle 52.

Notably, the motor heat exchanger 46 is provided upstream in fuel line 42 flow of the valve 54. Consequently, the motor heat exchanger 46 is provided between the fuel tank 44 and the valve 54 in fuel flow, and may be provided between the fuel tank and the oil-fuel heat exchanger 50.

The provision of the motor heat exchanger 46 between the fuel tank 44 and the fuel injectors 17 in fuel flow series has a number of advantages. In normal operation, while the internal combustion engine 10 and motor 28 are running, the fuel drain valve 52 is closed. Fuel flows through each of the heat exchangers 46, 50, to the fuel injectors 17, before being burned and the combustion products expelled.

In an emergency case, it may be necessary to dump fuel through the dump nozzle 52. This may be the case where for example, a failure has occurred shortly after take-off, and the aircraft is above a maximum safe landing weight, in view of the weight of the fuel. In such a case, the valve 54 will be opened, and fuel will flow out through the dump nozzle 52, lightening the aircraft. In view of the placement of the heat exchanger 46 between the fuel tank 44 and fuel drain valve 54, fuel will continue to flow past the heat exchanger 46 during fuel dumping. This is important in a hybrid aircraft, since it may be necessary to continue to operate the motor 28 at high power during such an emergency.

For example, one scenario is where the gas turbine engine 10 fails. In such a case, it may be necessary to dump fuel and land. Since the gas turbine engine 10 has failed, it will not be possible to burn fuel to lighten the aircraft, since fuel flow to the engine 10 will have ceased. On the other hand, the motor 28 may need to be operated at higher power, in view of the reduction of thrust that was previously provided by the gas turbine engine 10 prior to the failure. Consequently, the motor 28 may have high cooling requirements under such circumstances, and so it is important that coolant continues to flow under such conditions.

Under some circumstances, the arrangement of the fuel / cooling system of the present disclosure may have further advantages. For example, under normal operation, with both the gas turbine engine 10 and motor 28 being operational, the cooling demands for the motor 28 may be relatively low, in view of the relatively low power requirements, since thrust requirements are shared between the gas turbine engine 10 and motor 28. Consequently, a small motor heat exchanger 46 can be provided, which results in a low weight. Typically, electric motor power is limited by the cooling system, and so the motor 28 and cooling system are sized appropriately. In view of the relatively low flow of coolant (i.e. the fuel) under most conditions, only relatively low amounts of heat may be conducted to the fuel.

It should also be noted that fuel is used as a coolant for multiple aircraft systems. For example, the oil system requires cooling, and fuel is used for this purpose. The fuel has a limited heat capacity, and it is vital that the fuel is not heated beyond a certain temperature, known as the "coking temperature". At temperatures above the coking temperature, carbon deposits are formed in the fuel, which may block fuel injectors and result in decreased engine performance, and potential failure.

However, in the present case, where the engine 10 fails, these considerations no longer apply. The fuel temperature can be allowed to exceed the fuel coking temperature, since fuel no longer needs to be delivered to the engine 10. Furthermore, the drain valve 54 can be opened (since, in an emergency, range requirements are typically reduced, as the aircraft must only divert to a close-by airfield), thereby enabling increased fuel cooling flow. Consequently, in an emergency, fuel cooling capability is greatly increased, allowing for higher electric machine power. As a result, a system is provided in which high power density is provided, in view of the small cooling system and power dense electric motor.

Figure 4 shows an alternative aircraft 101. The aircraft 101 is similar to the aircraft 1, in that the aircraft comprises a fuselage 102, wings 103, tail 104 and propulsion system 105. However, the propulsion system 105 differs from the propulsion systems 5.

The propulsion system 105 comprises a plurality of propulsors in the form of fans 112a-e, each of which is coupled to a respective motor 128a-e.

In this embodiment, four fans 112a-d are provided mounted to the wings 103. The aircraft 101 further comprises a fifth propulsor 112e in the form of a boundary layer ingesting propulsor 112e mounted at the aft end of the fuselage 102, close to the tail 104. The fifth propulsor also comprises a fan, which is coupled to an electric motor 128e. By providing a propulsor at the aft end of the fuselage 102, boundary layer is ingested by the propulsor 112e, thereby increased propulsive efficiency of that propulsor, and so reducing overall aircraft thrust requirements.

The propulsion system 105 further comprises an internal combustion engine in the form of a gas turbine engine 110. The engine 110 is similar the engine 10, and is mechanically coupled to a generator 132. The generator 132 is configured to generate electrical power as the gas turbine engine 110 operates, and is electrically coupled to the electric motors 128a-e, and to an energy storage device in the form of a battery 130. Since the gas turbine engine 110 is not itself mechanically coupled to a propulsor, the propulsion system 105 can be said to be a "series hybrid" aircraft.

Figure 5 shows a schematic drawing of a fuel and cooling system of the propulsion system 105. A fuel tank is provided 144, which provides fuel to a fuel line 142. The fuel line carries fuel to injectors 117 of the gas turbine engine 10. A motor heat exchanger 146 is provided, which exchanges heat between fuel in the fuel line 142, and an intermediate coolant in intermediate coolant lines 148a-e (only two of which are shown in figure 5 for clarity. Second heat exchangers 148a-e (only two of which are shown in figure 5) are provided, which exchange heat between respective electric motors 128a-e, and the intermediate coolant. Consequently, the second heat exchangers 148a-e and motor heat exchanger 146 cooperate to cool the motors 128a-e, in a similar manner to the system shown in figure 3.

The remainder of the fuel / cooing system is similar to that of figure 3. Again, a fuel drain valve 154 and dump nozzle 152 are provided, which are each downstream of the motor heat exchanger 146. Similar, an oil to fuel heat exchanger 150 is provided between the motor heat exchanger 146 and fuel drain valve 154.

Again therefore, motor cooling can continue where the gas turbine engine is inoperative, such as where the engine 10 is shutdown in an emergency. Consequently, the advantages outlined above in relation to the parallel hybrid system also apply to some extent to the series hybrid system of figure 5.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

For example, gas turbine engines having more or fewer shafts could be provided. Electric motors having different topologies could be provided, such as topologies in which the stator is provided radially inside of the rotor. Different fuels and coolant fluids could be utilised. Different components of the motors could be cooled using the cooling system.

Claims

Claims 1 An aircraft propulsion system (5) comprising; an internal combustion engine (10); an electric propulsor motor (28); a propulsor (12) configured to be driven by at least the propulsor motor (28); wherein the electric motor (28) comprises an internal combustion engine fuel cooled cooling system (40) comprising a motor heat exchanger (46) configured to transfer heat from the electric motor (28) to internal combustion engine fuel, and wherein the propulsion system further comprises an overboard fuel dump system comprising an overboard drain valve (54) provided downstream in fuel flow of the motor heat exchanger (46).
2. A system according to claim 1, wherein the internal combustion engine (10) comprises one or more fuel injectors (17) downstream of the overboard drain valve (54).
3. A system according to claim 1 or claim 2, wherein the aircraft propulsion system (105) comprises a series hybrid propulsion system, in which the internal combustion engine (110) is coupled to an electric generator (132) configured to provide electrical power to the propulsor motor (128), and wherein the internal combustion engine (110) is not mechanically coupled to a propulsor (112, 152).
4. A system according to claim 1 or claim 2, wherein the aircraft propulsion system (5) comprises a parallel hybrid propulsion system, in which the internal combustion engine (10) is mechanically coupled to at least one propulsor (12).
5. A system according to claim 5, wherein the internal combustion engine (10) is coupled to an electric generator (32) configured to provide electrical power to the propulsor motor (28).
6. A system according to any of the preceding claims, wherein the aircraft propulsion system (5) comprises an energy storage device (30) configured to store electrical energy for provision to the propulsor motor (28).
7. A system according to any of the preceding claims, wherein the internal combustion engine comprise a gas turbine engine (10).
8. A system according to any of the preceding claims, wherein the internal combustion engine (10) comprises a cooling system configured to utilise internal combustion engine fuel as a heat sink.
9. A system according to claim 8, wherein the internal combustion engine cooling system comprises an oil cooling system, which comprises a fuel-oil heat exchanger (50) configured to reject heat from the internal combustion engine oil to the internal combustion engine fuel.
10.A system according to claim 9, wherein the fuel-oil heat exchanger (50) is provided downstream in fuel flow of the motor heat exchanger (46).
11.A system according to any of the preceding claims, wherein the motor cooling system comprises an intermediate cooling loop, comprising a second heat exchanger (48) configured to exchange heat between the electric motor (28) and an intermediate cooling fluid, the motor heat exchanger (46) being configured to exchange heat between the intermediate cooling fluid and the fuel.
12.A system according to any of the preceding claims, wherein the internal combustion engine fuel comprise a liquid fuel such as a liquid hydrocarbon.
13.An aircraft (1) comprising a propulsion system (5) according to any of the preceding claims.