Patent Description:
Aviation is one of the fastest-growing sources of greenhouse effect and gas emissions. However, environmental footprint commitments have to be met. Emissions, like CO2, NOx, SOx, other pollutants (e.g. PMx) and noise need to be reduced.

Thus, the thermal engine cycle based on hydrocarbon fuels needs to be replaced soon to reduce this problem. For regional travel, principal big actor of emission, a switch from conventional gas turbine engine to hybrid and/or electric aircraft has been identified as a growing market with a high number of aeroplanes in the next years and decades.

However, while electric propulsion has advanced very rapidly, battery technology hasn't. The main issue is that <NUM>% of the aviation industry's emissions come from passenger flights longer than <NUM> - a distance no electric airliner could yet fly with current battery technology.

An alternative solution is to use hydrogen fuel, which makes no CO2 emission if hydrogen is created in CO2 neutral manner.

<CIT> discloses a fuel cell system for an aircraft and a method for operating the fuel cell system in the aircraft. The fuel cell system comprises fuel cell units coupled to a hydrogen tank.

<CIT> discloses a method for the combustion of hydrogen in a burner inside a combustion chamber of a gas turbine, where hydrogen and air are injected into the burner.

<CIT> discloses an aircraft propulsion system with a hydrogen driven turbine and a fuel cell.

It is the object of the invention to reduce the environmental footprint of aircrafts by reducing their emissions.

The object is achieved by a hybrid propulsion system for aircraft, comprising at least a fuel cell unit for providing power to drive an aircraft by using hydrogen in a redox reaction, at least a gas turbine for providing power to drive the aircraft by burning hydrogen, a hydrogen source for providing the hydrogen to the fuel cell unit and/or to the gas turbine, and at least an electric propulsion unit for driving the aircraft, the electric propulsion unit comprising a first electric machine connected to a first propeller, the first electric machine being electrically powered by the fuel cell unit and mechanically powered by the gas turbine in at least two different propulsion modes, and a control unit to control the power supply to the electric propulsion unit, wherein the control unit is configured to apply a first propulsion mode for a first flight phase in which the electric propulsion unit is essentially or only electrically powered by the fuel cell unit, and to apply a second propulsion mode for a second flight phase in which the electric propulsion unit is essentially or only mechanically powered by the gas turbine.

Emissions like CO2, NOx, SOx, other pollutants (e.g. PMx) and noise are reduced. Moreover, <NUM>% emissions can be achieved e.g. in cities and airports, i.e. no pollution, no CO2, NOx, SOx, other pollutants are emitted and noise is reduced due to gas turbine inoperative. In addition, <NUM>% CO2 emissions can be achieved e.g. during cruise where only a residual small amount of NOx is emitted, but no SOx and no other pollutants. Thus, the pollution on the overall flying phase is reduced.

The invention creates an innovative Zero-C dual H2 aircraft concept which combines both, H2 Burn and fuel cell propulsion in different flight phases and thus reduces harmful emissions. H2 Burn uses hydrogen as a fuel for e.g. a conventional gas turbine engine, and fuel cell propulsion uses the hydrogen for a redox reaction on an anode-cathode cell.

The control unit may be configured to apply the first propulsion mode in particular in one or more of the flight phases taxiing, take-off, first part of climbing, second part of descent and landing. In this way, <NUM>% emissions are achieved in these phases, and there is no pollution in cities and airports, and no CO2, NOx, SOx, other pollutants are emitted and noise is reduced there.

The control unit may be configured to apply the second propulsion mode in particular in one or more of the flight phases second part of climbing, cruise and first part of descent. In this way, <NUM>% CO2 emissions and a highly reduced amount of NOx are achieved during cruise and no SOx and no other pollutants are emitted in these phases.

According to a preferred embodiment, the electric propulsion unit may comprise a second electric machine for driving a second propeller, wherein the fuel cell unit is adapted to provide the electrical power to the first electric machine and/or to the second electric machine, and wherein the gas turbine is adapted to provide the mechanical power at least to the first electric machine and to the first propeller via a common shaft.

Thus, the hybrid propulsion system is redundant. It enables also to still operate the aircraft in a safe manner in case of failure.

The first electric machine may be adapted to convert the mechanical energy provided by the gas turbine into electrical energy, and to provide that electrical energy e.g. to the second electric machine to drive its assigned propeller.

The control unit may comprise a failure control automatism for automatically changing from one to the other propulsion mode in case of a failure of an element of the propulsion chain, in case of a failure of the fuel cell unit and/or in case of a failure of an electric connection for delivering electrical energy to an electric machine of the electric propulsion unit, and/or in case of a failure of an electric machine of the electric propulsion unit, and/or in case of a failure of a gas turbine. Depending on the type of a failure and the currently activated propulsion mode, the control unit may activate a flow of hydrogen to the fuel cell unit and/or in addition to the gas turbine, and/or a flow of electrical power from the fuel cell unit to one or more of the electric machines for driving propellers, and/or in addition a flow of mechanical power from the gas turbine to both electric machines or to only one of the electric machines.

In the first propulsion mode, preferably only the fuel cell unit may be powered by the hydrogen. In the second propulsion mode, preferably only the gas turbine may be powered by the hydrogen.

According to an aspect of the invention, a method of operating a hybrid propulsion system of an aircraft is provided, the hybrid propulsion system comprising an electric propulsion unit which comprises a first electric machine connected to a first propeller and a gas turbine, wherein the method comprises the steps of providing hydrogen to a fuel cell unit and/or to a gas turbine, providing power to drive the electric propulsion unit of the aircraft in a first propulsion mode by using the hydrogen for a redox reaction in the fuel cell unit, providing power to mechanically drive the electric propulsion unit of the aircraft in a second propulsion mode by burning the hydrogen in the gas turbine, wherein in a first flight phase the first propulsion mode is essentially or only used, and in a second flight phase the second propulsion mode is essentially or only used.

The first propulsion mode may be applied in particular in one or more of the flight phases taxiing, take-off, first part of climbing, second part of descent and landing.

The second propulsion mode may be applied in particular in one or more of the flight phases second part of climbing, cruise and first part of descent.

For example, in the first propulsion mode only the fuel cell unit may be powered by the hydrogen.

For example, in the second propulsion mode only the gas turbine may be powered by the hydrogen.

Depending on a type of a failure and the currently activated propulsion mode, a flow of hydrogen to the fuel cell unit and/or at the same time to the gas turbine may be activated, and/or a flow of electrical power from the fuel cell unit and/or in addition a flow of mechanical power from the gas turbine to the first electric machine and/or to a second electric machine of the electric propulsion unit connected to a second propeller may be activated.

According to a preferred embodiment, in the first propulsion mode the fuel cell unit may provide electrical power to the first electric machine for driving the first propeller and/or to the second electric machine of the electric propulsion unit for driving the second propeller, and in the second propulsion mode the gas turbine may provide mechanical power at least to the first electric machine for driving the first propeller.

In the second propulsion mode the gas turbine may provide mechanical power to a first electric machine which may transform the mechanical power to provide electrical power to the second electric machine for driving another propeller.

In the second propulsion mode, electrical energy may be produced by the first electric machine which converts mechanical energy provided by the gas turbine and which may provide the electrical energy to the second electric machine for driving another propeller.

According to a further preferred embodiment, the method comprises a failure operation in which the current propulsion mode is replaced by the other propulsion mode in case of a failure of an element of the propulsion chain, in case of a failure of the fuel cell unit, and/or in case of a failure of an electric connection delivering electrical energy to the electric machine for driving the propeller, and/or in case of a failure of the electric machine of the electric propulsion unit for driving the aircraft, and/or in case of a failure of a gas turbine.

Advantageously, a hybrid propulsion system according to the invention can be used in the method according to the invention.

According to another aspect, the invention provides a hybrid aircraft which comprises a hybrid propulsion system according to the invention.

Characteristics and advantages described in relation to the hybrid propulsion system are also related to the method of operating a hybrid propulsion system and vice versa.

In the following, exemplary embodiments showing further advantages and characteristics are described in detail with reference to the figures, in which:.

In the figures, similar or identical elements and features are designated by the same reference numbers.

<FIG> shows a hybrid propulsion system <NUM> according to a first embodiment of the invention. The propulsion system <NUM> comprises a gas turbine <NUM> which is connected to a fuel tank <NUM> which contains hydrogen, whereby the hydrogen can be supplied via duct <NUM> to the gas turbine <NUM>. The gas turbine <NUM> is e.g. a conventional gas turbine for burning the hydrogen. In addition, a fuel cell unit <NUM> is connected to the fuel tank <NUM> via duct <NUM> in order to be supplied with hydrogen.

An electric propulsion unit <NUM> is mechanically connected to the gas turbine <NUM> and electrically connected to the fuel cell unit <NUM> whereby both can provide energy (mechanical and electrical) to the electric propulsion unit <NUM>, wherein the electric propulsion unit <NUM> is essentially powered by the fuel cell unit <NUM> in a first propulsion mode, whereas in a second propulsion mode the electric propulsion unit <NUM> is essentially powered by the gas turbine <NUM>.

A power control unit <NUM> controls the hydrogen flow from fuel tank <NUM> to gas turbine <NUM> and/or to fuel cell unit 13as well as the flow of electrical power E and mechanical power Pm to the electric propulsion unit <NUM>.

The electric propulsion unit <NUM> comprises at least an electric machine 15a which is connected to a propeller 16a and to the gas turbine <NUM> on a common shaft. The electric machine 15a receives in the first propulsion mode mechanical power Pm from the gas turbine <NUM> via a shaft 33a in order to drive the propeller 16a via a shaft 33b.

In a second propulsion mode the electric machine 15a receives electrical power E from the fuel cell unit <NUM> via an electrical connection or electric line <NUM> and provides it as mechanical power Pm to propeller 16a. Thus, redundancy is given. The hybrid propulsion system <NUM> realises a Zero-C dual H2 system concept.

It is also possible that the electric machine 15a receives mechanical power Pm from the gas turbine <NUM> and in addition electrical power E from the fuel cell unit <NUM>.

The electric machine 15a is fully integrated and comprises an inverter and controller with inherent safety features to control switching between the different operational modes.

<FIG> shows the gas turbine <NUM> as an enlarged view, which is known in the state of the art. It receives hydrogen from the fuel tank <NUM>, which is indicated in the figure by arrow H<NUM>. On its air inlet at its front side the gas turbine <NUM> receives oxygen from the ambient air, which is indicated in the figure by arrow <NUM><NUM>. During operation, the gas turbine <NUM> burns the hydrogen, whereby energy is set free and water and NOx is produced. The energy is transformed in mechanical power and provided to the propeller <NUM>.

The gas turbine <NUM> produces <NUM>% CO2 and works with high efficiency while it produces high power.

In <FIG> the fuel cell or fuel cell unit <NUM> which is part of the Hybrid propulsion system <NUM>, is shown in an enlarged view. It is known in the state of the art and comprises an anode 13a, a cathode 13b and an electrolyte 13c which is disposed between the anode 13a and the cathode 13b (see <FIG>). During operation, the fuel cell <NUM> receives hydrogen at a fuel input opening 17a on its anode side, as indicated by arrow H2, and it receives oxygen at an oxidant input opening 17b on its cathode side, as indicated by arrow <NUM> (see <FIG>). As a redox reaction on the anode-cathode cell, and electric power is produced which can be supplied to an electric load <NUM> (see <FIG>).

The fuel cell <NUM> produces <NUM>% CO2 and works with relatively low efficiency and high power. A relatively big heat exchanger is foreseen.

Depleted fuel is output at a depleted fuel output opening 18a at the anode side of the fuel cell <NUM>. Depleted oxidant and H<NUM>O is output at a depleted oxidant and H<NUM>O output opening 18b at the cathode side of fuel cell <NUM>. By the redox reaction electricity e- is output from the fuel cell <NUM> as indicated by arrow E, and provided to the electric machine 15a which forms an electric motor and drives the propeller <NUM> (see <FIG>).

<FIG> shows the configuration and the operation of the hybrid propulsion system <NUM> depending on different flight phases of an aircraft <NUM>.

During flight phases taxiing, take-off and first part of climbing the hybrid propulsion system <NUM> is mainly or exclusively operated in a first mode M1, in which the power for taxiing, take-off and first part of climbing is solely or at least mainly generated by the fuel cell unit <NUM>. During taxiing a relatively low power is needed, whereas during take-off the power reaches a maximum. In this example, the phases taxiing and take-off have a duration of approximately <NUM> minutes.

During the flight phases second part of climbing, cruise and first part of descent the hybrid propulsion system <NUM> is mainly or exclusively operated in a second mode M2, in which the power for second part of climbing, cruise and first part of descent is solely or at least mainly generated by H2 burning in the gas turbine <NUM>. During climbing, the needed power is essentially lower than during take-off. During cruise the needed power is reduced, and during descent the power is further reduced. In this example, the phases second part of climbing, cruise and first part of descent have a duration of approximately <NUM> minutes.

During the second part of descent and landing phase the hybrid propulsion system <NUM> is again operated in the first mode M1 were the power is solely or at least mainly generated by the fuel cell unit <NUM>. The power level is lower than during cruise. In this example, the second part of descent and landing phase as a duration of approximately <NUM> minutes.

<FIG> shows a hybrid propulsion system <NUM> according to a second embodiment of the invention. As in the first embodiment shown in <FIG>, the propulsion system <NUM> comprises gas turbine <NUM>, fuel tank <NUM>, fuel cell unit <NUM> and an electric propulsion unit <NUM>. Gas turbine <NUM>, fuel tank <NUM> and fuel cell unit <NUM> are described above in detail with reference to <FIG>.

In this embodiment the electric propulsion unit <NUM> comprises in addition to the first electric machine 15a connected to first propeller 16a and to the gas turbine <NUM>, a second electric machine 15b connected to a second propeller 16b. In addition to the above description with reference to <FIG>, the first electric machine is also capable to convert the mechanical energy provided by the gas turbine <NUM> into electrical energy, and to provide that electrical energy to the second electric machine 15b to drive the second propeller 16b. The second electric machine 15b and the second propeller 16b have the same characteristics as the first electric machine 15a and the first propeller 16a, which are described above in detail with reference to <FIG>.

The fuel tank <NUM> containing hydrogen is connected with the fuel cell unit <NUM> and the gas turbine <NUM> by ducts <NUM>, <NUM>. The fuel cell unit <NUM> is connected electrically to the electric machines 15a, 15b of the electric propulsion unit <NUM> by electric lines <NUM>, <NUM>. Power control unit <NUM> controls valves <NUM>, <NUM> arranged in the ducts <NUM>, <NUM> respectively to control the H2 flow to the fuel cell unit <NUM> and to the gas turbine <NUM> depending on the operational phase or flight phase of the aircraft <NUM> as well as the electric and mechanical power transmission through the electric lines <NUM>, <NUM> and shafts 33a, 33b depending on the operational phase or flight phase of the aircraft <NUM>. The power control unit <NUM> thus provides an H2, electric and mechanical energy management within the hybrid propulsion system.

A method of operating the hybrid propulsion system is shown in the following with reference to <FIG>, <FIG> and <FIG>, where different flow paths are described.

Energy flow path <NUM> is applied during taxiing, take-off, first part of climbing, second part of descent and landing phase of aircraft <NUM>. H2 flows from fuel tank <NUM> through ducts <NUM> to fuel cell <NUM>. Electric energy flows from the fuel cell <NUM> through electric lines <NUM>, <NUM> to the first and second electric machine 15a, 15b. Mechanical power Pm is transmitted from electric machines 15a, 15b through shafts 33b, 33c to propellers 16a, 16b. This describes the first propulsion mode M1 which is applied by the power control unit <NUM> during the period of these flight phases.

In propulsion mode M1 indicated by first path <NUM>, the H2 powers the fuel cell <NUM> only and by consequence both electric machines 15a, 15b. The gas turbine <NUM> is not powered by H2. Being controlled by the power control unit <NUM>, the valve <NUM> which connects the H2 tank <NUM> with the fuel cell unit <NUM> is in an open state, whereas the valve <NUM> which connects the H2 tank <NUM> with the gas turbine <NUM> is in closed state.

By the redox reaction in fuel cell unit <NUM>, electric energy is output from the fuel cell <NUM> and delivered to the electric machines 15a, 15b. There, the electric energy is converted into mechanical power Pm and transmitted to propellers 16a, 16b.

When the second part of the climbing phase begins, energy flow path <NUM> is activated by the power control unit and will automatically apply the second propulsion mode M2.

In this mode, the H2 will power the gas turbine <NUM> only and by consequence the associated propeller 16a via shaft 33a, 33b through the shaft of the first electric machine 15a. In addition for <FIG> and <FIG>, the gas turbine <NUM> provides energy to the second electric machine 15b via the first electric machine 15a, which serves as a generator, and via electric lines <NUM>, <NUM> electrically connecting first and second electric machines 15a, 15b. The second electric machine 15b serves as a motor for driving second propeller 16b.

The fuel cell unit <NUM> will not be powered by H2, i.e. it will not supply power to the electric propulsion unit <NUM> in mode M2.

To achieve this, valve <NUM> is controlled by the power control unit <NUM> to be in an open state and thus allows an H2 flow from the H2 tank <NUM> to the gas turbine <NUM> through duct <NUM>, whereas valve <NUM> is controlled to be in a closed state whereby no H2 is delivered to the fuel cell unit <NUM>.

The propulsion mode M2 is still applied in the first part of the descent phase in this example of operating the hybrid propulsion system <NUM>. When it comes to the second part of the descent phase, the system will commute back to propulsion mode M1 in which only the fuel cell unit <NUM> is powered and flow path <NUM> is activated.

The configuration and the operation of the hybrid propulsion system <NUM> depending on the flight phases is shown in the following with reference to <FIG> and <FIG> according to the second embodiment.

<FIG> shows the different propulsion modes M1 and M2 depending on the flight phases, and <FIG> show the configuration of the hybrid propulsion system <NUM> in the different phases according to the second embodiment.

<FIG> is similar to <FIG> which is described above, but further shows the propulsion power in MW (megawatts).

<FIG> shows the configuration of the propulsion system <NUM> during propulsion mode M1 applied during taxiing, take-off and first part of climbing according to the second embodiment. The valve <NUM> is in an open state and valve <NUM> is closed whereby H2 flows through duct <NUM> to the fuel cell unit <NUM> only. In this example, <NUM> MW are provided on each side of the aircraft <NUM> in the e.g. take-off phase. The fuel cell unit <NUM> powered by the H2 supply delivers electrical power E to both electric machines or motors, which then provide mechanical power Pm in equal shares of <NUM> MW to the propellers 16a, 16b. Thus, the total power on both sides of the aircraft <NUM> is <NUM> MW in this example.

<FIG> shows the configuration of the propulsion system <NUM> during propulsion mode M2 applied during second part of climbing, cruise and first part of descent. Here, valve <NUM> is in an open state whereas valve <NUM> is closed whereby H2 flows to the gas turbine <NUM> only. In this example, <NUM> MW are provided on each side of the aircraft <NUM> in e.g. the climbing phase.

The gas turbine <NUM> powered by the H2 supply delivers mechanical power PM to first electric machine 15a which acts as a generator to generate a flow of electric power E to second electric machine 15b, while it also supplies mechanical power Pm to drive its associated propeller 16a. The power is delivered to both propellers 16a, 16b in equal shares of <NUM> MW. Thus, the total power on both sides of the aircraft <NUM> is <NUM> MW in this example.

<FIG> shows the configuration of the propulsion system <NUM> during propulsion mode M1 again, to which the system commutes at the beginning of the second part of the descent phase. The configuration of the propulsion system <NUM> including the H2 flow and electric and mechanic power flows is similar to the configuration described above with reference to <FIG>, however it generates less power.

<FIG> show different failure mode operations of the hybrid propulsion system <NUM> during e.g. take-off, which are described in the following. As shown below, the propulsion system <NUM> is redundant and enables still to operate in a safe manner the aircraft <NUM> in case of a failure.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during a fuel cell unit <NUM> failure in e.g. the take-off phase, in which the fuel cell <NUM> is inoperative or shows a malfunction. The system switches automatically to the H2 burning configuration. In this configuration the second propulsion mode M2 is applied. H2 is supplied from fuel tank <NUM> to the gas turbine <NUM> only. Valve <NUM> is closed and valve <NUM> is open.

The gas turbine <NUM> provides <NUM> MW to the first electric machine 15a, which forwards half of the power as electrical power E to the second electric machine 15b, and half of the power as mechanical power Pm to its associated propeller 16a. Thus, both propellers 16a, 16b on one side of the aircraft <NUM> are operative and each of them provides <NUM> MW. In total, the propulsion energy on both sides of the aircraft <NUM> amounts to <NUM> MW.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during an electric connection failure in e.g. the take-off phase. In this case the electric connection provided by connection lines <NUM>, <NUM> is interrupted. The system <NUM> activates automatically the H2 burning configuration similar to second propulsion mode M2. The gas turbine <NUM> supplies <NUM> MW. In addition, the fuel cell <NUM> provides <NUM> MW.

H2 is supplied from fuel tank <NUM> to the fuel cell unit <NUM> and in addition to the gas turbine <NUM>. Both valves <NUM>, <NUM> are open and both propellers 16a, 16b on one side of the aircraft <NUM> are operative and each of them provides <NUM> MW. In total, the propulsion energy on both sides of the aircraft <NUM> amounts to <NUM> MW.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during an electric machine failure in e.g. the take-off phase.

In case that the external second electric machine 15b becomes inoperative, the system <NUM> automatically activates the H2 burning automatic configuration, similar to the second operational mode M2. The gas turbine <NUM> provides <NUM> MW. In addition, the fuel cell <NUM> provides <NUM> MW.

H2 is supplied from fuel tank <NUM> to the fuel cell unit <NUM> and in addition to the gas turbine <NUM>. Both valves <NUM>, <NUM> are open. Gas turbine <NUM> supplies power to its associated first electric machine 15a, and in addition fuel cell unit <NUM> supplies via electric lines <NUM>, <NUM> electrical power E to the first electric machine 15a. Thus, the first propeller 16a is powered with <NUM> MW while the external second propeller 16b is not operative.

In case that the internal first electric machine 15a is inoperative, the system <NUM> automatically activates the H2 burning automatic configuration, similar to the second operational mode M2. In addition, the fuel cell <NUM> further generates power.

H2 is supplied from fuel tank <NUM> to the fuel cell unit <NUM> which supplies <NUM> MW in this example, and in addition to the gas turbine <NUM> which also supplies <NUM> MW in this example. Both valves <NUM>, <NUM> are open and both propellers 16a, 16b are operative.

<FIG> show different failure mode operations of the hybrid propulsion system <NUM> during e.g. cruise, which are described in the following.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during a turbine failure in the e.g. cruising phase, in which the turbine <NUM> is inoperative or shows a malfunction. The system switches automatically to the fuel cell automatic configuration. In this configuration the first propulsion mode M1 is applied. H2 is supplied from fuel tank <NUM> to the fuel cell unit <NUM> only. Valve <NUM> is closed and valve <NUM> is open.

The fuel cell <NUM> provides <NUM> MW, which is distributed by electric connections <NUM>, <NUM> to both electric machines 15a, 15b whereby both propellers 16a, 16b are operative and each of them provide <NUM>. In total, the propulsion energy on both sides of the aircraft <NUM> amounts to <NUM> MW.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during an electric connection failure during e.g. cruise. In this case the electric connection provided by connection lines <NUM>, <NUM> is interrupted. The system <NUM> activates automatically the fuel cell configuration which is similar to propulsion mode M1. In addition, the gas turbine <NUM> generates power.

H2 is supplied from fuel tank <NUM> to fuel cell unit <NUM>, which provides <NUM> MW in this example, and in addition to the gas turbine <NUM> which also provides <NUM> MW. Both valves <NUM>, <NUM> are open. Both propellers 16a, 16b on each side of the aircraft <NUM> are operative and each of them provides <NUM> MW.

<FIG> shows the configuration of the hybrid propulsion system <NUM> during an electric machine failure during e.g. cruise. In case that the external electric machine 15b is inoperative, the system <NUM> automatically activates the fuel cell automatic configuration similar to the first operational mode M1. The fuel cell provides power, in this example <NUM> MW. In addition, the gas turbine <NUM> provides power, in this example also <NUM> MW.

H2 is supplied from fuel tank <NUM> to the fuel cell unit <NUM> and in addition to the gas turbine <NUM>. Both valves <NUM>, <NUM> are open. Fuel cell unit <NUM> supplies electrical power E to the internal first propeller 16a which also receives the mechanical power Pm from gas turbine <NUM>. Thus, the internal first propeller 16a is operative while the external second propeller 16b is not operative.

In case that the internal first electric machine 15a is inoperative, the system <NUM> automatically activates the fuel cell automatic configuration similar to the first operational mode M1. In addition, the gas turbine <NUM> provides power.

H2 is supplied from fuel tank <NUM> to the gas turbine <NUM> which supplies <NUM> MW in this example, and in addition to the fuel cell unit <NUM> which also supplies <NUM> MW in this example. Both valves <NUM>, <NUM> are open and both propellers 16a, 16b are operative.

<FIG> shows a schematic view of an aircraft <NUM> equipped with the hybrid propulsion system <NUM> according to the first embodiment of the invention. In this example, a non-redundant dual-H2 concept architecture is realized.

Depending on the flying phase, the system will either commute to fuel cell propulsion or to H2 burn propulsion. This is achieved by two different flow paths <NUM> and <NUM>.

Path <NUM> is activated during at least one of the phases taxiing, take-off, first part of climbing, second part of descent and landing, preferably in all of these phases. The H2 supplied by fuel tank <NUM> will power the fuel cell <NUM> only and by consequence electric machine 15a associated therewith. The gas turbine <NUM> will not be powered. The propulsion system is operated in a way that valve <NUM> provided in duct <NUM> is closed while valve <NUM> provided in duct <NUM> is open. The fuel cell <NUM> provides electrical power E to the electric machine 15a which acts as a motor and which is connected to propeller 16a.

Path <NUM> is activated during the second part of climbing, cruise phase and first part of descent. It can also be activated during the climbing phase, preferably in both of these phases. The H2 supplied by fuel tank <NUM> will power the gas turbine <NUM> only and by consequence the propeller 16a. The fuel cell <NUM> will not be powered. The propulsion system is operated in a way that valve <NUM> provided in duct <NUM> is open and valve <NUM> provided in duct <NUM> is closed.

Summarized, the invention provides a dual-H2 concept architecture which can be redundant according to particularly preferred embodiments. It combines an H2-burn gas turbine with a fuel cell and at least an electric propulsion unit in a hybrid propulsion system. A fully integrated electric machine (motor/generator) including inverter/controller with inherent safety features can be part of the hybrid propulsion system. In particular, a pitch control management can be provided. In addition, a high power density technology with full redundancy can be implemented.

Further, the hybrid propulsion system may comprise an electric transmission system including e.g. a new cable technology, a voltage compatible with integrated and lightweight power electronics as well as high voltage protection devices.

With respect to control, monitoring and protection, one or more of the features power management (multi-source/multi-consumer), automatic reconfiguration upon failure detection as well as health monitoring can be provided.

With respect to thermal management or cooling system, a fuel cell thermal management, an electric propulsion unit thermal management and/or an H2 tank thermal management can be provided in the hybrid propulsion system.

Claim 1:
Hybrid propulsion system (<NUM>, <NUM>) for aircraft, comprising
at least a fuel cell unit (<NUM>) for providing power to drive an aircraft (<NUM>) by using hydrogen in a redox reaction;
at least a gas turbine (<NUM>) for providing power to drive the aircraft (<NUM>) by burning hydrogen;
a hydrogen source (<NUM>) for providing the hydrogen to the fuel cell unit (<NUM>) and/or to the gas turbine (<NUM>); and at least
an electric propulsion unit (<NUM>, <NUM>) for driving the aircraft (<NUM>), the electric propulsion unit (<NUM>, <NUM>) comprising a first electric machine (15a, 15b) connected to a first propeller (16a, 16b), the first electric machine being electrically powered by the fuel cell unit (<NUM>) and mechanically powered by the gas turbine (<NUM>) in at least two different propulsion modes (M1, M2); and
a control unit (<NUM>) to control the power supply to the electric propulsion unit (<NUM>, <NUM>),
wherein the control unit (<NUM>) is configured to apply a first propulsion mode (M1) for a first flight phase in which the electric propulsion unit (<NUM>, <NUM>) is essentially or only electrically powered by the fuel cell unit (<NUM>), and to apply a second propulsion mode (M2) for a second flight phase in which the electric propulsion unit (<NUM>, <NUM>) is essentially or only mechanically powered by the gas turbine (<NUM>).