HYBRID AIRCRAFT

Description

[0001] The invention relates to the field of aircraft and more particularly to the field of electrically powered aircraft with vertical takeoff and landing.

[0002] The aeronautics industry is currently undergoing significant changes, partly due to evolving environmental requirements and partly due to the development of electrically powered aircraft. In particular, the VTOL (Vertical Take-Off and Landing) sector is especially dynamic, offering very promising prospects as a new mode of transportation.

[0003] VTOL aircraft are a relatively old field in themselves (they were developed as early as 1921), but their electrification has led to a proliferation of new solutions and regulations. 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 in all propulsion and flight-related systems, from engines and power sources to the entire electrical system, to ensure continuous safe flight and landing (known as "continued safe flight and landing") and not just an emergency landing following a failure. This is also called "one-fail-safe," meaning "tolerant of a single failure."

[0004] These regulations pose numerous problems, particularly regarding feasibility while maintaining a reasonable cost. Indeed, if all components are doubled, then costs more than double because oversizing is necessary to manage the added weight, not to mention that it also means revising all the flight capabilities of the aircraft, which is thus heavier.

[0005] Therefore, other solutions must be found. Most solutions are based on two principles, possibly used together: Highly sophisticated energy management at the battery pack level, with a dedicated module that precisely controls their operating point. US patent 9,586,690 describes this type of solution. The use of a separate power generation source to power horizontal thrusters, thus only drawing power from the batteries during takeoff and landing, thereby increasing the range and endurance of electric propulsion systems. WO 2020 / 016510 and EP 3,628,593 describe this type of solution.

[0006] However, these solutions are not entirely satisfactory, and in particular do not provide a one-fail-safe solution.

[0007] US patent 6,293,491 describes a vertical takeoff and landing (VTOL) aircraft with a plurality of lift and thrust rotors, in which each rotor has its own electric motor for drive but without interconnection between drive groups. EP patent 3,296,212 describes a multi-switch architecture with horizontal-drive thrusters only. FR patent 3,095,806 describes a VTOL aircraft in which the various drive groups cannot be interconnected. None of these patents provides a one-fail-safe solution.

[0008] The invention improves the situation. To this end, it proposes a hybrid-powered aircraft that includes: at least two horizontal drive thrusters, each powered by an electric motor, forming a horizontal drive group; at least four pairs of vertical takeoff / landing rotors, each powered by an electric motor; and at least four stored electrical energy sources, each connected to a respective electric motor of a pair of vertical takeoff / landing rotors, forming with the corresponding electric motors and stored electrical energy source a vertical drive group comprising a power supply bus whose output can be connected to a single horizontal drive group, which can be connected to at least two vertical drive groups via a switch; at least two electrical generation sources connected on one side to each of the power supply buses by a respective input of the corresponding vertical drive group.and on the other hand, to each horizontal drive group via the respective output of each vertical drive group, at least one power supply control arranged to issue a power command to the power generation sources according to the power requirements of the vertical drive groups and / or the horizontal drive groups, the stored electrical energy sources supplying electricity according to the difference between the power requirements of the vertical drive groups and / or the horizontal drive groups and the power emitted by the power generation sources based on the power command, the power generation sources being further equipped to recharge the stored electrical energy sources, so that the stored electrical energy sources are passively controlled.

[0009] This aircraft is particularly advantageous because its architecture allows for redundancy, guaranteeing one-fail-safe operation while optimizing component sizing. Thus, the aircraft according to the invention minimizes the additional costs associated with implementing one-fail-safe operation and employs a truly hybrid architecture in which the power sources are genuinely complementary at each stage of flight.

[0010] According to various embodiments, the invention may have one or more of the following characteristics: The stored electrical energy sources for the vertical drive units are batteries; the fuel-electric power generator is a turbine generator; and the current converters are AC-to-DC converters. The fuel-electric power generator uses a fuel-based fuel, biofuel, or synthetic fuel. The fuel-electric power generator is a hydrogen fuel cell, and the current converters are DC-to-DC converters. The aircraft further includes electrical contactors connecting each of the horizontal drive units, the vertical drive units, and the power generation source to the rest of the aircraft's electrical system. Each element of each of the horizontal drive units,The vertical drive groups and the electrical power source are connected to the other elements of the horizontal drive group, the vertical drive group, or the electrical power source by an electrical contactor. The aircraft further includes diodes at the input and / or output of each of the horizontal drive groups, the vertical drive groups, and the electrical power source, connecting them to the electrical buses that link them. The diodes are arranged upstream of the contactors relative to the electrical power source.

[0011] Other features and advantages of the invention will become clearer upon reading the following description, drawn from illustrative and non-limiting examples taken from the drawings shown: there figure 1 represents a schematic view of the electrical architecture of an aircraft according to the invention, the figure 2 represents an example of a flight sequence indicating which elements are active in which phase and what the associated electrical charge levels are, the figure 3 represents an example of a configuration in case of loss of a vertical drive group during takeoff or landing, the figure 4 represents an example of a configuration in case of loss of an electrical generation source, the figure 5 represents an example of a configuration in case of loss of a horizontal drive group, the figure 6 represents an example of a configuration in case of loss of the electrical generation source, and the figure 7 represents an example of a power control management algorithm implemented by the aircraft.

[0012] The drawings and description below contain, for the most part, elements of a definite nature. They can therefore not only serve to better explain the present invention, but also contribute to its definition, if necessary.

[0013] This description contains material that may be protected by copyright. The rights holder has no objection to the reproduction of this patent document or its description, as it appears in official records, by anyone. Otherwise, the rights holder reserves all rights.

[0014] As can be seen on the Figure 1 , an aircraft 2 according to the invention comprises a control 4, two horizontal drive groups 6 and 8, four vertical drive groups 10, 12, 14 and 16, and an electrical generation source 18.

[0015] In the example described here, the horizontal drive unit 6 (or 8) comprises a DC-to-AC converter 22 (or 32), an electric motor 24 (or 34), and a propeller 26 (or 36), for example, a propeller. The propeller 26 (or 36) is arranged to enable the aircraft to move in a substantially horizontal direction. In the example described here, the propeller 26 (or 36) consumes 80 kW of power during flight.

[0016] The horizontal drive group 6 (respectively 8) is connected at the input to a switch 28 (respectively 38) which allows this input to be connected to the output of the vertical drive group 10 (respectively 14) or 12 (respectively 16), as described below.

[0017] The vertical drive unit 10 (respectively 12, 14, 16) comprises a rotor 42 (respectively 46, 72, 76) driven by a motor 52 (respectively 56, 82, 86), and a rotor 44 (respectively 48, 74, 78) driven by a motor 54 (respectively 58, 84, 88). The motors 52 and 54 are powered by a DC-to-AC converter 62 and 64 (respectively 66 and 68, 92 and 94, 96 and 98). DC to AC converters 62 and 64 (respectively 66 and 68, 92 and 94, 96 and 98) are connected to an electrical bus of the vertical drive group 10 (respectively 12, 14, 16), to which is connected a battery 50 (respectively 60, 80, 90) as well as a first input connected to a first electrical distribution bus of the electrical generation source 18 and a second input connected to a second electrical distribution bus of the electrical generation source 18.Finally, the electrical bus of each of the vertical drive groups 10 and 12 (respectively 14 and 16) is connected to a respective output of these groups, which is connected to switch 28 (respectively 38). As will be seen below, the batteries 50, 60, 80, and 90 each form a source of stored electrical energy and together deliver 600 kW when operating at 100% of their capacity. Alternatively, the batteries 50, 60, 80, and 90 could be supercapacitors or a combination of batteries and supercapacitors.

[0018] In the example described here, the power generation source 18 comprises a turbine generator 100 and two AC-to-DC converters 104 and 106. As explained in the preceding paragraph, the AC-to-DC converters 104 and 106 are each connected to a respective input of the power bus of the vertical drive unit 10 (12, 14, 16, respectively). Thus, the AC-to-DC converter 104 defines the starting point of the first power distribution bus of the power generation source 18, and the AC-to-DC converter 106 defines the starting point of the second power distribution bus of the power generation source 18.

[0019] In the example described here, the turbine generator can deliver 80 kW at 100% of its capacity. Alternatively, the power generation source could be different, using direct or alternating current followed by an AC-to-DC converter or a DC-to-DC converter. This source could be based on a turbogenerator powered by conventional fuel, biofuel, or synthetic fuels. Alternatively, a hydrogen-based energy source, such as a fuel cell, could be used. For the purposes of this invention, the turbine generator, the turbogenerator, and the hydrogen-based energy source are considered fuel-powered electricity generators.

[0020] As will be seen below, control 4 is a low-voltage device arranged to control, on the one hand, the electrical generation source 18, and on the other hand, the switches 28 and 38, as well as various protective elements not shown in the diagram. Figure 1 which will be explained further with the figures 2 à 6 .

[0021] When we analyze the Figure 1 It appears that all the motor and electrical components are redundant. Thus, one-fail-safe operation can be ensured, as will be described later. Indeed, there are two horizontal drive groups and four vertical drive groups, themselves forming two subgroups connected to a single horizontal drive group. As will be seen below, the power generation source also features redundancy via current converters.

[0022] But beyond this fairly classic duplication, it is the electric buses specific to each vertical drive group, as well as the electric distribution bus specific to each source of electrical generation, that make it possible to obtain the benefits of the invention together.

[0023] Indeed, as we will see below, the aircraft's particular structure Figure 1 This allows for true hybridization of electrical energy sources, as opposed to existing solutions where they are simply juxtaposed. Thus, depending on power requirements, both the batteries and the electrical generation source can operate in tandem. But beyond that, this architecture allows the batteries to be treated as pure "energy buffers." The batteries are managed entirely passively, without any need for software or hardware intelligence other than the basic intelligence required to operate the battery management system (BMS) itself, for example, to activate protections and report status.This goes completely against all existing solutions, in which either an element is specifically designed to optimize battery operation and plays a control role, or an element is designed to compensate for a possible weakness in the batteries, but in exclusive alternation, that is to say without the batteries and this element being likely to operate simultaneously.

[0024] There Figure 2 represents an energy consumption cycle during a flight with an aircraft of the Figure 1 As can be seen in this figure, the flight begins with an initial 200-meter maneuver in which the aircraft takes off vertically. During this phase, the rotors of the vertical drive units operate and consume the majority of the energy. The horizontal drive units may also operate for stability purposes, but with negligible energy consumption. They are powered by the batteries (600 kW) and the electrical generator (80 kW). Thus, the batteries, which are between 75% and 90% charged at the start of the flight, drop to between 55% and 70% of their capacity. At the end of this phase, the aircraft is approximately 50 feet from its takeoff point, representing a vertical drop of about 15 meters.

[0025] Next, in a 210 maneuver, the aircraft gradually transitions from vertical to horizontal flight at an altitude of between 50 and 150 feet, then performs a climb similar to that of conventional airplanes. During this phase, the rotors are progressively shut down, and power consumption decreases from 680 kW to 80 kW once horizontal cruising is achieved. The batteries and the electrical power source continue to operate at full capacity, and the batteries continue to discharge, losing 10% to 30% of their charge by the time horizontal cruising is reached.

[0026] The horizontal flight is performed in a 220 operation during which the batteries are not used. The electrical power source continues to run at full capacity, and the 80 kW it produces is shared between the horizontal drive units, which are controlled by the flight controller. The power not consumed by these units is used to recharge the batteries. This phase, at an altitude of over 1,000 feet (approximately 300 m), recharges the batteries to about 50%.

[0027] The descent then proceeds in a 230 operation, in which the electrical power source is used at 100% to recharge the batteries. This allows the batteries to be fully recharged. During this phase, no motor components consume energy.

[0028] The transition from horizontal to vertical flight is then carried out in a 240 operation, in which consumption gradually increases to approximately 340kW, and in which the battery charge gradually decreases from 100% to between 85% and 95%.

[0029] Finally, in operation 250, the vertical landing is performed using only the rotors, as in takeoff, but with the advantage of using gravity. Thus, the batteries continue to discharge to between 75% and 90%, as at the starting point of operation 200.

[0030] It therefore appears that the aircraft does not need to be recharged on the ground between flights, which increases its usability. Furthermore, it is clear that the electrical power generation source always operates at full capacity (there is one exception that will occur with the Figure 7 ), and that the batteries serve to compensate for situations where the electrical power source cannot provide enough power. Similarly, as soon as possible, the batteries are recharged as much as possible to ensure sufficient electrical energy is available for landing.

[0031] Alternatively, the aircraft's batteries can be recharged on the ground between two flights sufficiently separated in time. In this case, control 4 can make more sophisticated trade-offs regarding power supply during the various phases, either to increase the operating points of the turbine generator to modify the distribution of energy supply and increase the range, limit noise, pollutant emissions, etc.

[0032] Battery control is managed passively thanks to the architecture of the invention - if the rotors or horizontal drive groups draw less than 80kW, then the batteries are naturally not used, and the surplus current can even be used to recharge them (as in the 220 or 230 operation); if more power is required, then the batteries are naturally used.

[0033] It also appears that the oversizing of the batteries can be kept quite low thanks to the architecture of the invention. Indeed, it is no coincidence that between 10% and 30% of the charge remains in the batteries at the end of the 220 operation – this guarantees that the one-fail-safe condition can be maintained with a single failure wiping out one battery.

[0034] There Figure 3 This represents a case of failure of a battery or other electrical component of one of the vertical drive units. For simplicity, only the energy output from the electrical generation source 18 and the vertical drive unit 12 is shown, but the other components function similarly.

[0035] In the case shown, the first vertical drive group 10 is taken out of service due to an electrical failure of one of the motors 52 or 54. This is particularly disabling during one of the operations 200, 210, 240 or 250.

[0036] First, it should be noted that each electrical component 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) at the inputs connected to the first and second electrical distribution buses of the power generation source 18 allow the vertical drive unit 10 (respectively 12, 14, 16) to be isolated and prevent any propagation of an electrical problem from the outside to the vertical drive unit 10 (respectively 12, 14, 16). This is also the case for the horizontal drive units with switches 28 and 38, as well as for the power generation source with switches (not shown).In addition, each element within these electrical subsets is also connected to the rest of the electrical subset to which it belongs by means of an unshown switch, so that, if for example battery 50 malfunctions, it can be cut off from the rest of the vertical drive group 10 without immediately disconnecting the latter.

[0037] In the example described here, the contactors are doubled by the presence of diodes (not shown) which allow, in the event of an electrical problem and in particular a short circuit, the vertical drive group to be isolated passively to prevent any propagation of electrical problem from the vertical drive group 10 (respectively 12, 14, 16) to the outside.

[0038] Furthermore, batteries 50, 60, 80, and 90 are oversized. Simultaneously, control 4 ensures that switch 28 connects to the output of the vertical drive unit 12. The batteries are then subjected to a load of 100 kW. In addition, the power generation unit can be operated beyond its normal operating point, reaching 110% or 120% for a few minutes. This, along with the increased load on the batteries, compensates for the 80 kW loss due to the disconnection of the vertical drive unit 10. The path of the electricity leaving the power generation unit 18 is shown by bold arrows. Thus, takeoff (or an emergency landing) is guaranteed and can be performed safely, at the cost of slightly oversizing the batteries.

[0039] There Figure 4 represents a different failure case in which the AC to DC converter 104 is lost.

[0040] As can be seen in this figure, the inputs of the vertical drive groups connected to the electrical distribution bus of the AC to DC converter 104 are isolated by means of contactors opened by control 4. In addition, the electrical generation source is stressed beyond its classical operating point, and batteries are used in order to maintain a power supply of around 70kW, i.e. 90% of the power normally consumed by the horizontal drive groups.

[0041] Here again, one-fail-safe is guaranteed, since the batteries can be recharged during the descent phase before landing. Furthermore, the 25% larger battery capacity ensures sufficient energy to continue horizontal flight at 90% of normal capacity.

[0042] There Figure 5 This illustrates a different failure scenario in which a horizontal drive unit is lost. In this case, the remaining horizontal drive unit is used to its maximum capacity to maintain horizontal cruising with an altitude and speed profile that accounts for the loss. The batteries can be used as a buffer in case of sudden power surges.

[0043] There Figure 6 This illustrates yet another different failure scenario in which the fuel-powered power generator fails. In this case, the first and second distribution power buses of power generation source 18 are deactivated, and batteries 50, 60, 80, and 90 provide the necessary power to ensure flight continuity and guarantee one-fail-safe operation.

[0044] Indeed, the loss of the power generation source is compensated by the stored electrical energy sources, which constitutes a safety gain through dissimilar redundancy (i.e., replacing one source with another of a different type). This provides an additional level of safety compared to simple redundancy of the power generation source. However, this safety through dissimilarity is asymmetrical in the sense that the loss of the power generation source can be compensated by the stored electrical energy sources, but conversely, the redundancy of the stored energy sources is used to ensure one-fail-safe operation.

[0045] In this case of failure, two scenarios are considered: either the control 4 detects that the voltage in battery 50 (respectively 80) and in battery 60 (respectively 90) are sufficiently close, in which case the distribution group including contactors not shown at the inputs connected to the first electrical distribution bus of the electrical generation source 18 (respectively to the second electrical distribution bus of the electrical generation source 18) connects the two batteries 50 and 60 (respectively 80 and 90) to the horizontal drive group 6 (respectively 8) to smooth the energy consumption on the two batteries simultaneously, or the control 4 detects that the voltages are not close and the switch 28 (respectively 38) is first connected to one of the two batteries 50 and 60 (respectively 80 and 90) until it is discharged, then is connected to the other of the two batteries.

[0046] There Figure 7shows an algorithm that can be implemented by command 4 to manage the power control according to the various moments.

[0047] This is a cycle that starts in an operation 700 from the reception of an operating point for the rotors and / or the thrusters.

[0048] Next, in operation 710, the rotors and / or thrusters draw a current corresponding to that operating point. This operation is followed by a test in operation 720 to determine whether the turbine generator is operating at 100% capacity. If it is not, then control 4 pushes it to full capacity, and while it increases its speed, the batteries compensate for the current demand in operation 730. If the turbine generator is operating at full capacity, then in operation 740, control 4 determines whether its output is sufficient to meet the current demand of operation 710. If not, then full capacity is maintained until the next operating point, and the batteries are charged. If they are sufficient, then in operation 750, control 4 checks whether the batteries need to be charged.If so, then full load is maintained and the excess power is used to charge the batteries to the next operating point. Otherwise, then control 4 reduces the operating point of the turbine generator in a 760 operation, to the next operating point.

[0049] The power consumption figures above were given for illustrative purposes only and are not exhaustive. The electrical architecture should be adapted to the actual flight requirements of the aircraft.

[0050] It appears from the above that the stored electrical energy sources are of the high-power / low-capacity type, while the electrical generation source is of the high-capacity / low-power type. This follows from the fact that the stored electrical energy sources are used as energy buffers and that the electrical generation source is sized to meet the power consumption during horizontal flight and to allow for the recharging of the energy buffers.

[0051] This duality, implemented through the architecture described above in which all electrical circuits are interconnected yet made independent by protective devices, creates redundancy that guarantees one-fail-safe operation while optimizing component sizing. Thus, the aircraft according to the invention minimizes the additional costs associated with implementing one-fail-safe operation and employs a truly hybrid architecture in which the power sources are genuinely complementary at each stage of flight.

[0052] It should also be noted that the figures represent electrical diagrams of the aircraft. Thus, although rotors 42 and 44, 46 and 48, 52 and 54, and 56 and 58 are shown side by side, this will not necessarily be the case from a mechanical point of view. Indeed, the rotors are assembled in pairs in a vertical drive unit so that a failure does not destabilize the aircraft. The rotors in the same vertical drive unit will therefore generally be arranged symmetrically with respect to the center of the aircraft.

[0053] Furthermore, although the figures depict an aircraft with 2 propellers, 8 rotors, and 2 turbine generators, their actual numbers may differ. Indeed, there may be more than 2 propellers, and there may be more than two vertical drive units per horizontal drive unit. Similarly, control 4 may also be duplicated to provide an additional degree of resilience.

[0054] Finally, the fact that a given vertical drive unit can only be connected to one horizontal drive unit simplifies the control of the architecture. This principle allows for control of redundancy and sizing by limiting the number of different vertical drive units connected to the same horizontal drive unit. Resilience is managed by switching the horizontal drive units, which can be implemented easily. This is much more efficient than a scheme where a single vertical drive unit could be connected to multiple horizontal drive units, which would create significant control problems in both steady-state and degraded operating conditions.

[0055] Alternatively, as suggested above, the aircraft could be ground-charged so that the stored electrical energy sources are 100% full at takeoff. This would allow for the implementation of other flight patterns, increasing the aircraft's range, as well as limiting noise and pollutant emissions during low-altitude takeoffs and landings.

Claims

1. Aircraft having a hybrid power source comprising: - at least two horizontal drive thrusters (26, 36), each powered by a respective electric motor (24, 34), forming at least two respective horizontal drive units (6, 8), - at least four pairs of vertical take-off / 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 sources of stored electrical power (50, 60, 80, 90), each connected to a respective electric motor (52, 54, 56, 58, 82, 84, 86, 88) of a pair of vertical take-off and landing rotors (42, 44, 46, 48, 72, 74, 76, 78), each pair of rotors (42, 44, 46, 48, 72, 74, 76, 78) forming with the respective electric motors (52, 54, 56, 58, 82, 84, 86, 88) and the corresponding source of stored electrical power a vertical drive unit (10, 12, 14, 16), each vertical drive unit (10, 12, 14, 16) comprising an electrical power supply bus, the outlet 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 a respective switch (28, 38) disposed at the inlet of the horizontal drive unit (6, 8), - a single electrical power generation source (18) comprising a fuel-based electrical power generator and at least two current converters (104, 106) each connected on the one hand to each of the electrical power supply buses via a respective inlet of the corresponding vertical drive unit (10, 12, 14, 16), and on the other hand to each horizontal drive unit (6, 8) via the respective outlet of each vertical drive unit (10, 12, 14, 16), - at least one electrical power supply controller (4) arranged to transmit a power command to the electrical power generation source (18) according to the power requirements of the vertical drive units (10, 12, 14, 16) and / or of the horizontal drive units (6, 8), the sources of stored electrical power (50, 60, 80, 90) supplying electricity according to the difference between the power requirements of the vertical drive units (10, 12, 14, 16) and / or of the horizontal drive units (6, 8) and the power output by the electrical power generation source (18) based on the power command, the electrial power generation source (18) being further capable of recharging the sources of stored electrical power (50, 60, 80, 90), such that the sources of stored electrical power (50, 60, 80, 90) are passively controlled.

2. Aircraft according to claim 1, wherein the sources of stored electrical power (50, 60, 80, 90) of the vertical drive units (10, 12, 14, 16) are batteries.

3. Aircraft according to claim 1 or 2, wherein the fuel-based electrical power generator is a turbine generator (100) and the current converters are AC to DC converters (104, 106).

4. Aircraft according to claim 3, wherein the fuel-based electrical power generator uses a fuel based on motor fuel, biofuel, or synthetic fuel.

5. Aircraft according to claim 1 or 2, wherein the fuel-based electrical power generator is a hydrogen fuel cell (100) and the current converters are DC to DC converters (104, 106).

6. Aircraft according to one of the preceding claims, further comprising electrical contactors connecting each of the horizontal drive units (6, 8), of the vertical drive units (10, 12, 14, 16) and the electrical power generation source (18) to the rest of the electrical circuit of the aircraft.

7. Aircraft according to claim 6, wherein each element of each of the horizontal drive units (6, 8), of the vertical drive units (10, 12, 14, 16) and of the electrical power generation source (18) is connected to the other elements of the respective horizontal drive unit (6, 8), of the vertical drive unit (10, 12, 14, 16) or of the electrical power generation source (18) via an electrical contactor.

8. Aircraft according to one of the preceding claims, further comprising diodes at the inlet and / or at the outlet of each of the horizontal drive units (6, 8), of the vertical drive units (10, 12, 14, 16) and of the electrical power generation source (18) connecting them to the electrical buses connecting them.

9. Aircraft according to claims 6 or 7 taken in combination with claim 8, wherein the diodes are disposed upstream of the contactors relative to the electrical power generation source (18).