Aircraft with a Hybrid Power Supply

Abstract

An aircraft having a hybrid power source includes at least two horizontal drive propellers forming a horizontal drive unit, each powered by an electric motor; at least four pairs of vertical takeoff and landing rotors, each powered by an electric motor; and at least four storage-type electric energy sources respectively connected to the electric motors of each pair of vertical takeoff and landing rotors. Together with the electric motors and the corresponding storage-type electric energy sources, a vertical drive unit including a power supply bus is formed. The output of the power supply bus can be connected to a single horizontal drive unit that can be connected to at least two vertical drive units via switches. The at least four pairs of vertical takeoff and landing rotors and the at least four storage-type electric energy sources are respectively connected to each of the power supply buses by the respective inputs of the corresponding vertical drive units and are connected to each horizontal drive unit via the respective outputs of each vertical drive unit. At least two power generation sources are arranged to issue power commands to the power generation sources according to the power requirements of the vertical drive units and / or the horizontal drive units. The storage-type electric energy sources supply electricity according to the difference between the power requirements of the vertical drive units and / or the horizontal drive units and the power discharged by the power generation sources based on the power commands. The power generation sources are further adapted to charge the storage-type electric energy sources, whereby the storage-type electric energy sources are passively controlled, including at least one power supply controller.

Description

Technical Field

[0001]

[0001] The present invention relates to the field of aircraft, and more particularly to the field of electric propulsion vertical take-off and landing aircraft.

Background Art

[0002]

[0002] The aviation industry is currently experiencing many changes, which are related in part to the evolution of environmental requirements and in part to the development of electric propulsion aircraft. In particular, the field of VTOL (representing vertical take-off and landing) is particularly active because it shows very interesting prospects as a new means of transportation.

[0003]

[0003] VTOL itself is a rather old field (developed as early as 1921), but with their electrification, the number of new solutions and regulations provided has increased explosively. In particular, the latest regulations (see, for example, SC-VTOL-01 SPECIAL CONDITION Vertical Take-Off and Landing (VTOL) Aircraft issued on July 2, 2019) require redundancy of all systems, engines, power sources, and electrical systems related to propulsion and flight in order to ensure not only emergency landing after a failure but also flight continuity (referred to as "continuous safe flight and landing"). This is also called "one-fail-safe", that is, "resistance to one failure".

[0004]

[0004] These regulations have many problems in terms of being achievable while maintaining a particularly reasonable cost. In fact, doubling all elements will result in a cost increase of more than double because it is necessary to increase the size to cope with the additional weight, and needless to say, this requires a review of the overall flight capabilities of the aircraft with such increased weight.

[0005]

[0005] Therefore, it is necessary to find other solutions. Most solutions are based on the following two principles, which may be used in combination: - Highly advanced energy management at the battery pack level by a dedicated module that manages their operating points in detail. U.S. Patent No. 9,586,690 describes this type of solution. - Use of a power source that powers a horizontal thruster to use the battery only during takeoff and landing. This improves the range and durability of the electrically propelled solution. International Publication No. 2020 / 016510 and European Patent No. 3 628 593 describe this type of solution.

[0006]

[0006] However, these solutions are not entirely satisfactory and do not particularly enable the achievement of a one-fail-safe solution.

Summary of the Invention

Means for Solving the Problems

[0007]

[0007] The present invention improves this situation. For this purpose, the present invention is an aircraft having a hybrid power source, - At least two horizontal drive thrusters each powered by an electric motor, the at least two horizontal drive thrusters forming a horizontal drive unit, - At least four pairs of vertical takeoff and landing rotors each powered by an electric motor and at least four accumulative electric energy sources each connected to the electric motor of each pair of vertical takeoff and landing rotors, the electric motor and the corresponding accumulative electric energy source together forming a vertical drive unit including a power supply bus, and an output portion of the power supply bus can be connected to a single horizontal drive unit that can be connected to at least two vertical drive units via a switch, at least four pairs of vertical takeoff and landing rotors and at least four accumulative electric energy sources, - A single power generation source (18), comprising a fuel-based generator and at least two power converters, wherein on the one hand, each is connected to a respective power supply bus by an input of a corresponding vertical drive unit, and on the other hand, at least two power converters each connected to a respective horizontal drive unit via an output of a corresponding vertical drive unit, the single power generation source (18), - At least one power supply controller arranged to issue a power command to the power generation source in response to the power requirements of the vertical drive unit and / or the horizontal drive unit, wherein the storage-type electrical energy source supplies electricity according to the difference between the power requirements of the vertical drive unit and / or the horizontal drive unit and the power released by the power generation source based on the power command, and the power generation source is further adapted to charge the storage-type electrical energy source, whereby the storage-type electrical energy source is passively controlled, the at least one power supply controller providing an aircraft comprising.

[0008]

[0008] This aircraft is particularly advantageous because its architecture can achieve redundancy that guarantees one-fail-safe while optimizing the size of the components. Thus, the aircraft according to the present invention implements a true hybrid architecture with minimal additional costs associated with the implementation of one-fail-safe and where the power sources are truly complementary at each stage of flight.

[0009]

[0009] According to various embodiments, the present invention may have one or more of the following features: - The storage-type electrical energy source of the vertical drive unit is a battery, - The power generation source is a turbine generator and the power converter is an AC-DC converter, - The power generation source uses a combustible material based on fuel, biofuel or synthetic fuel, - The power generation source is a hydrogen fuel cell and the power converter is a DC-DC converter, - The aircraft further includes an electrical contactor that connects each of the horizontal drive unit, the vertical drive unit, and the power generation source to the rest of the aircraft's electrical circuit. - Each element of the horizontal drive unit, the vertical drive unit, and the power generation source is connected by an electrical contactor to other elements of the horizontal drive unit, the vertical drive unit, or the power generation source. - The aircraft further includes diodes at the input and / or output of each of the horizontal drive unit, the vertical drive unit, and the power generation source that connect them to an electrical bus connecting the horizontal drive unit, the vertical drive unit, and the power generation source. - The diodes are arranged upstream of the contactor with respect to the power generation source.

[0010]

[0010] Other features and advantages of the present invention will become more apparent upon reading the following description. The following description is obtained from examples given for the purpose of explanation and is not intended to be limiting, and is obtained from the following drawings.

Brief Description of the Drawings

[0011]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Figure 7

Modes for Carrying Out the Invention

[0012]

[0011] The drawings and the following description essentially include specific elements. Therefore, they are not only used to better understand the present invention, but may also contribute to its definition as necessary.

[0013]

[0012] This description may include elements that are subject to protection by the author's rights and / or copyrights. The rights owner does not object to any person reproducing the patent document or its description as set forth in the public file. In other cases, the rights owner reserves its rights in full.

[0014]

[0013] As shown in FIG. 1, the aircraft 2 according to the present invention includes a controller 4, two horizontal drive units 6 and 8, four vertical drive units 10, 12, 14, and 16, and a power generation source 18.

[0015]

[0014] In the example described in this specification, the horizontal drive unit 6 (or 8) includes a DC-AC converter 22 (or 32), an electric motor 24 (or 34), and a propulsion unit 26 (or 36) having, for example, a propeller. The propulsion unit 26 (or 36) is arranged so that the aircraft can move forward substantially horizontally. In the example described in this specification, the propulsion unit 26 (or 36) consumes 80 kW of power during the flight mode.

[0016]

[0015] The horizontal drive unit 6 (or 8) is connected at its input to a switch 28 (or 38) that enables this input to be connected to the output of the vertical drive unit 10 (or 14) or 12 (or 16) as described later.

[0017]

[0016] The vertical drive unit 10 (or 12, 14, 16) includes a rotor 42 (or 46, 72, 76) driven by a motor 52 (or 56, 82, 86) and a rotor 44 (or 48, 74, 78) driven by a motor 54 (or 58, 84, 88). The motors 52 and 54 are powered by respective DC-AC converters 62 and 64 (or 66 and 68, 92 and 94, 96 and 98). The DC-AC converters 62 and 64 (or 66 and 68, 92 and 94, 96 and 98) are connected to the electrical bus of the vertical drive unit 10 (or 12, 14, 16), and this electrical bus is connected to a battery 50 (or 60, 80, 90) and a first input connected to the first electrical distribution bus of the power source 18 and a second input connected to the second electrical distribution bus of the power source 18. Finally, the respective electrical buses of the vertical drive units 10 and 12 (or 14 and 16) are connected to their respective output portions, and the output portions are connected to the switches 28 (or 38). As will be understood below, each of the batteries 50, 60, 80, and 90 forms an energy storage type electrical energy source and outputs 600 kW in total when outputting 100% of their capacities. Alternatively, the batteries 50, 60, 80, and 90 may consist of supercapacitors or a combination of batteries and supercapacitors.

[0018]

[0017] In the example described herein, the power source 18 includes a turbine generator 100 and two AC-DC converters 104 and 106. As explained in the previous paragraph, the AC-DC converters 104 and 106 are respectively connected to the respective input portions of the electrical buses of the vertical drive unit 10 (or 12, 14, 16). Therefore, the AC-DC converter 104 defines the starting point of the first electrical distribution bus of the power source 18, and the AC-DC converter 106 defines the starting point of the second electrical distribution bus of the power source 18.

[0019]

[0018] In the examples described herein, the turbine generator can output 80 kW at 100% of its capacity. Alternatively, the power source can be different and can be DC or AC followed by an AC-DC converter or a DC-DC converter. Thus, this power source can be based on a turbo generator powered by conventional fuel, biofuel or synthetic fuel. Further alternatively, a hydrogen-based energy source such as a fuel cell can also be used. In the context of the present invention, the turbine generator 100, the turbo generator, and the hydrogen-based energy source are considered to be fuel-based power generation devices.

[0020]

[0019] As will be understood hereinafter, the controller 4 is a low-voltage device arranged to control, on the one hand, the power sources 18 and 20 and, on the other hand, the switches 28 and 38, and also controls various protection elements (not shown in FIG. 1 but further described in FIGS. 2-6).

[0021]

[0020] Analyzing FIG. 1, it can be seen that all motors and electrical elements are duplicated. This allows for ensuring one-fail-safe as will be described later. In fact, there are two horizontal drive units and four vertical drive units formed by themselves in two subgroups connected to the same horizontal drive unit. As will be understood hereinafter, the power source also features duplication via a current inverter.

[0022]

[0021] However, beyond this very conventional duplication, the electrical buses specific to each vertical drive unit and the electrical distribution buses specific to each power source together make it possible to obtain the advantages of the present invention.

[0023]

[0022] In fact, as will be understood below, the specific structure of the aircraft of FIG. 1 enables true hybridization of the electrical energy source, which is in contrast to the existing solutions with which it is juxtaposed. Thus, depending on the power requirements, both the battery and the power generation source can operate in concert. However, not only that, this architecture enables the battery to be treated simply as an "energy buffer". The battery is treated completely passively and requires no software or hardware intelligence other than the basic intelligence necessary to operate the BMS (Battery Management System) of the battery system itself, for example to activate protection functions and report status. This is completely contrary to all existing solutions, where an element is specially made to optimize the operation of the battery, functioning as a controller, or an element is made to compensate for potential weaknesses of the battery. However, it is an exclusive alternative, i.e., the battery and this element cannot operate simultaneously.

[0024]

[0023] FIG. 2 shows the energy consumption cycle during flight of the aircraft of FIG. 1. As shown in this figure, the flight begins with a first operation 200 in which the aircraft takes off vertically. At this stage, the rotor of the vertical drive unit operates and accounts for most of the energy consumed. The horizontal drive unit may operate to ensure stability, but the energy consumption is minimal. They are supplied with up to 600 kW of power from the battery and up to 80 kW of power from the power generation source. Thus, the battery, which was charged between 75% and 90% at the start of the flight, shifts to a capacity of between 55% and 70%. At the end of this stage, the aircraft is at approximately 50 feet from the takeoff point, i.e., there is an altitude difference of approximately 15 meters.

[0025]

[0024] Subsequently, in operation 210, the aircraft gradually transitions from vertical flight to horizontal flight at 50 feet to 150 feet, and then ascends in the same manner as a conventional aircraft. At this stage, the rotor gradually stops, and when it reaches horizontal cruise, the power consumption drops from 680 kW to 80 kW. The battery and the power generation source continue to operate at full capacity, the battery continues to discharge, and when it reaches the horizontal cruise stage, it drops to 10% - 30%.

[0026]

[0025] Horizontal flight is executed in operation 220. During this period, the battery is not used. The power generation sources continue to operate at full capacity, and the 80 kW they generate is distributed among the horizontal drive units controlled by the controller, and the power not consumed by them is used to charge the battery. This stage exceeds an altitude of 1,000 feet (about 300 m) and enables the battery to be charged to about 50%.

[0027]

[0026] Subsequently, descent occurs in operation 230, and the power generation source is used at 100% to charge the battery. As a result, the battery is charged to 100%. At this stage, the motor elements do not consume energy.

[0028]

[0027] Subsequently, the transition from horizontal flight to vertical flight is carried out in operation 240, and the power consumption gradually transitions to about 340 kW, and the battery charge also gradually transitions from 100% to 85% - 95%.

[0029]

[0028] Finally, in operation 250, vertical landing is carried out using only the rotor as in takeoff, but taking advantage of gravity. Therefore, the battery continues to discharge to 75 - 90% as at the start of operation 200.

[0030]

[0029] Thus, the aircraft does not need to be charged on the ground between two flights, thereby improving its usability. Furthermore, it is clear that the power generation source is always operating at full capacity (with the exception cases shown in FIG. 7), and the battery serves to compensate when the power generation source cannot supply sufficient power. Similarly, as quickly as possible, the battery is charged as much as possible to ensure sufficient electrical energy for landing.

[0031]

[0030] Alternatively, the aircraft battery can be charged on the ground between two flights with sufficient time intervals. In this case, the controller 4 can raise the operating point of the turbogenerator to change the power supply distribution, extend the flight range, and perform more advanced trade-offs regarding power supply during various stages, such as limiting noise pollution and pollutant emissions.

[0032]

[0031] The control of the battery is passively managed by the architecture of the present invention. When the rotor or the horizontal drive unit draws less than 80 kW of power, naturally the battery is not used, and the surplus current can also be used for charging the battery (such as in operations 220 or 230). When more power is required, naturally the battery is used.

[0033]

[0032] The architecture of the present invention can also significantly suppress the enlargement of the battery. In fact, it is not accidental that 10% - 30% of the charge remains in the battery at the end of operation 220, which guarantees that a one-fail-safe can be maintained even if a failure occurs in the battery.

[0034]

[0033] FIG. 3 shows an example of a failure of the battery or other electrical elements in one of the vertical drive units. For the sake of brevity, only the energy from the power generation source 18 and the vertical drive unit 12 is shown, but other elements operate similarly.

[0035]

[0034] In the illustrated example, due to an electrical failure of either motor 52 or 54, the first vertical drive unit 10 is switched off. This particularly causes problems during one of operations 200, 210, 240, or 250.

[0036]

[0035] First, it should be noted that each electrical element is protected by a contactor that can be controlled to isolate it from the rest of the circuit. Thus, in this embodiment, two contactors (not shown) of the input section connected to the first and second electrical distribution buses of the power source 18 isolate the vertical drive unit 10 (or 12, 14, 16) and prevent the propagation of electrical problems from the outside to the vertical drive unit 10 (or 12, 14, 16). This also applies to the horizontal drive units with switches 28 and 38 and the power source with switches not shown. Further, each element within these electrical subsets is also connected via a switch (not shown) to the rest of the electrical subset to which it belongs, so that, for example, if the battery 50 malfunctions, it can be disconnected from the rest of the vertical drive unit 10 without being immediately disconnected.

[0037]

[0036] In the example described herein, the contactor is duplicated by the presence of a diode (not shown), and the diode can isolate the vertical drive unit in a passive manner in the event of an electrical problem, particularly a short circuit, and prevent the propagation of electrical problems from the vertical drive unit 10 (or 12, 14, 16) to the outside.

[0038]

[0037] Further, the batteries 50, 60, 80 and 90 are oversized. At the same time, by the controller 4, the switch 28 is surely connected to the output part of the vertical drive unit 12. Therefore, the batteries are used up to 100 kW. Further, the power generation source can be used beyond the conventional operating point and can be used at 110% or 120% for several minutes. Thus, combined with the excessive use of the batteries, the 80 kW loss related to the disconnection of the vertical drive unit 10 is compensated. The path that the electricity from the power generation source 18 follows is indicated by the thick arrow.

[0039]

[0038] Thus, takeoff (or emergency landing) is guaranteed and can be carried out without risk by slightly increasing the size of the batteries.

[0040]

[0039] FIG. 4 shows different failure cases when the AC-DC converter 104 is lost.

[0041]

[0040] As shown in this figure, the input part of the vertical drive unit connected to the electrical distribution bus of the AC-DC converter 104 is isolated by a contactor opened by the controller 4. Further, the power generation source is used beyond the conventional operating point, and the batteries are taken in to maintain power supply in the range of 70 kW. This corresponds to 90% of the power normally consumed by the horizontal drive unit.

[0042]

[0041] Also in this case, since the batteries can be charged during the descent stage before landing, one-fail-safe is guaranteed. Further, by making the batteries 25% larger, it is guaranteed to hold enough energy to continue horizontal flight at 90% of the normal capacity.

[0043]

[0042] FIG. 5 shows yet another different failure case, and this time the lost one is the horizontal drive unit. In this case, the remaining horizontal drive units are utilized to the maximum of their capabilities and horizontal cruising is maintained with an altitude and speed profile considering the loss. The batteries can be used as a buffer in case of sudden excessive consumption.

[0044]

[0043] FIG. 6 also shows different failure cases, in which the fuel-based generator does not function properly. In this case, the first electrical distribution bus of the power source 18 and the second electrical distribution bus of the power source 18 stop functioning, and the batteries 50, 60, 80, and 90 supply the energy necessary to ensure the continuation of the flight and guarantee a one-fail-safe.

[0045]

[0044] In fact, the loss of the power source is compensated by the storage-type electrical energy source, which is an improvement in heterogeneous redundancy safety (i.e., one supply source is replaced by another supply source of a different type). This consists of a further level of safety compared to the simple redundancy of the power source. On the other hand, the safety due to this heterogeneity is asymmetric, and the loss of the power source can be compensated by the storage-type electrical energy source. In the reverse case, the redundancy of the stored energy source is used to guarantee a one-fail-safe.

[0046]

[0045] In this failure case, two scenarios are conceivable: - The controller 4 detects that the voltages of the batteries 50 (or 80) and the batteries 60 (or 90) are sufficiently close. In this case, in the input connected to the first electrical distribution bus of the power source 18 (or the second electrical distribution bus of the power source 18), the distribution group including the contactor (not shown) connects the two batteries 50 and 60 (or 80 and 90) to the horizontal drive unit 6 (or 8) and simultaneously smooths the energy consumption in the two batteries, or - When the controller 4 detects that the voltages are not close, the switch 28 (or 38) is first connected to one of the two batteries 50 and 60 (or 80 and 90), and when it becomes empty, it is connected to the other of the two batteries.

[0047]

[0046] FIG. 7 shows the algorithm executed by the controller 4 to manage the power commands according to various situations.

[0048]

[0047] This consists of a cycle that starts with an operation 700 that receives the operating point of the rotor and / or the propeller.

[0049]

[0048] Then, in operation 710, the rotor and / or the propeller draws the current corresponding to this operating point. Following this operation, in operation 720, a test is performed to determine whether the turbogenerator is controlled at 100%. If it is not controlled at 100%, the controller 4 pushes it fully, and while the speed is increasing, in operation 730, the battery compensates for the current demand. When the turbogenerator is pushed fully, in operation 740, the controller 4 determines whether its power generation is sufficient for the current draw in operation 710. If it is not sufficient, the full load is maintained until the next operating point, and the battery is loaded. If it is sufficient, in operation 750, the controller 4 checks whether the battery needs to be charged. If necessary, the full load is maintained, and the surplus power is used to charge the battery until the next operating point. If not necessary, the controller 4 reduces the operating point of the turbogenerator in operation 760 until the next operating point.

[0050]

[0049] In the above, the power consumption is shown purely as an indicator and is not limiting. It is necessary to adapt the electrical architecture according to the actual requirements related to the flight of the aircraft.

[0051]

[0050] As described above, the storage-type electrical energy source is of the high-output / low-capacity type, and the power generation source is of the high-capacity / low-output type. This is because the storage-type electrical energy source is used as an energy buffer, the power generation source is sized to match the consumption during horizontal flight, and it enables the charging of the energy buffer.

[0052]

[0051] This duality is achieved by the above-described architecture in which all electrical circuits are interconnected independently by protection devices, enabling redundancy that ensures fail-safe while optimizing the size of the elements. Thus, the aircraft according to the present invention implements a true hybrid architecture with minimal additional cost associated with fail-safe implementation and in which the power sources are truly complementary at each stage of flight.

[0053]

[0052] Note also that the figure shows the electrical circuit diagram of the aircraft. Thus, although rotors 42 and 44, 46 and 48, 52 and 54, 56 and 58 are shown side by side, this is not necessarily the case from a mechanical point of view. In fact, the rotors are assembled in pairs with a vertical drive unit so that a failure does not destabilize the aircraft. Thus, the rotors of the same vertical drive unit are generally arranged symmetrically with respect to the center of the aircraft.

[0054]

[0053] Furthermore, the figure shows an aircraft having two propulsion engines, eight rotors, and two turbogenerators, but their numbers can be different. In fact, there may be more than three propulsion engines, and there may be more than three vertical drive units for one horizontal drive unit. Similarly, doubling command 4 can also guarantee additional restoring force.

[0055]

[0054] Finally, the fact that a given vertical drive unit can be connected to only one horizontal drive unit simplifies the control of the architecture. In fact, this principle makes it possible to control redundancy and sizing according to the number of different vertical drive units connected to the same horizontal drive unit. The restoring force is also organized by the switch of the horizontal drive unit, which can be implemented in a simple way. This is much more efficient than a scheme in which one vertical drive unit can be connected to multiple horizontal drive units, which can cause major control problems in both steady state and degradation mode.

[0056] Alternatively, as proposed above, the aircraft can be charged on the ground so that the power source is 100% full at takeoff. This increases the aircraft's range and enables the implementation of other flight schemes by limiting noise and pollutant emissions during low-altitude takeoff and landing phases.

Claims

Claim 1 An aircraft having a hybrid power source, - At least two horizontal drive propulsion units (26, 36) each powered by a respective electric motor (24, 34), the at least two horizontal drive propulsion units (26, 36) forming at least two respective horizontal drive units (6, 8), - At least four pairs of vertical takeoff and landing rotors (42, 44, 46, 48, 72, 74, 76, 78) each powered by a respective electric motor (52, 54, 56, 58, 82, 84, 86, 88) and at least four accumulative electric energy sources (50, 60, 80, 90) each connected to a respective electric motor (52, 54, 56, 58, 82, 84, 86, 88) of a pair of vertical takeoff and landing rotors (42, 44, 46, 48, 72, 74, 76, 78), each pair of rotors (42, 44, 46, 48, 72, 74, 76, 78) forming a vertical drive unit (10, 12, 14, 16) together with the respective electric motor (52, 54, 56, 58, 82, 84, 86, 88) and the corresponding accumulative electric energy source, each vertical drive unit (10, 12, 14, 16) including a power supply bus, the output of which can be connected to a single horizontal drive unit (6, 8), the number of vertical drive units (10, 12, 14, 16) being such that each horizontal drive unit (6, 8) can be connected to at least two vertical drive units (10, 12, 14, 16) via respective switches (28, 38) arranged at the input of the horizontal drive units (6, 8), at least four pairs of vertical takeoff and landing rotors (42, 44, 46, 48, 72, 74, 76, 78) and at least four accumulative electric energy sources (50, 60, 80, 90), - A single power generation source (18), comprising a fuel-based generator, and at least two power converters (104, 106), each of which is connected to a respective one of the power supply buses by a respective input of the corresponding vertical drive unit (10, 12, 14, 16) on the one hand, and on the other hand is connected to each horizontal drive unit (6, 8) via a respective output of each vertical drive unit (10, 12, 14, 16). - At least one power supply controller (4) arranged to issue a power command to the power generation source (18) in response to the power requirements of the vertical drive units (10, 12, 14, 16) and / or the horizontal drive units (6, 8), wherein the storage-type electrical energy source (50, 60, 80, 90) supplies electricity in response to the difference between the power requirements of the vertical drive units (10, 12, 14, 16) and / or the horizontal drive units (6, 8) and the power discharged by the power generation source (18) based on the power command, and the power generation source (18) is further adapted to charge the storage-type electrical energy source (50, 60, 80, 90), whereby the storage-type electrical energy source (50, 60, 80, 90) is passively controlled. At least one power supply controller (4). An aircraft comprising the same.

2. The aircraft according to claim 1, wherein the storage-type electrical energy source (50, 60, 80, 90) of the vertical drive units (10, 12, 14, 16) is a battery.

3. The aircraft according to claim 1 or 2, wherein the power generation source is a turbine generator (100), and the power converters are AC-DC converters (104, 106).

4. The aircraft according to claim 3, wherein the power generation source uses a combustible material based on fuel, biofuel or synthetic fuel.

5. The power generation source is a hydrogen fuel cell (100), and the power converter is a DC-DC converter (104, 106). The aircraft according to claim 1 or 2.

6. The aircraft according to any one of claims 1 to 5, further comprising an electrical contactor that connects each of the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), and the power generation source (18) to the rest of the aircraft's electrical circuit.

7. For the aircraft according to claim 6, each element of each of the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), and the power generation source (18) is connected by an electrical contactor to another element of the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), or the power generation source (18).

8. The aircraft according to any one of claims 1 to 7, further comprising a diode that connects the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), and the power generation source (18) to the electrical bus that connects them at each input and / or output of each of the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), and the power generation source (18).

9. The aircraft according to claim 6 or 7, considered in combination with claim 8, wherein the diode is arranged upstream of the contactor with respect to the power generation source (18).

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