Hybrid-powered aircraft

Abstract

The present invention relates to an aircraft with a hybrid power supply. The aircraft comprises two horizontal drive units (6, 24, 26), four vertical drive units (10, 42, 44, 52, 54, 50) and a vertical drive unit (10) including a power supply bus. An output of the power supply bus may be connected via a switch (28) to a single horizontal drive unit (6) which may be connected to at least two vertical drive units (10). The aircraft also comprises two power generation sources (18) connected on the one hand to each power supply bus via an input of the corresponding vertical drive unit (10) and on the other hand to each horizontal drive unit (6) via a corresponding output of each vertical drive unit (10). The aircraft also comprises a power supply control (4) arranged to send power commands to the power generation source (18) according to the power requirements of the vertical drive units (10) and / or the horizontal drive units (6). The power generation source (18) is suitable for recharging the power sources (50) of the vertical drive units.

Description

[Technical Field]

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

[0002] The aviation industry is currently undergoing a number of upheavals, partly linked to changing environmental requirements and the development of electric aircraft. The field of VTOL (Vertical Take-Off and Landing) is particularly dynamic, offering very interesting prospects for new modes of mobility.

[0003] While VTOL itself is a fairly old field (development began as early as 1921), its electrification has led to an explosion of new solutions and regulations. In particular, the latest regulations (e.g., "SC-VTOL-01 SPECIAL CONDITION Vertical Take-Off and Landing (VTOL) Aircraft," issued on July 2, 2019) require redundancy of all systems related to propulsion and flight, from engines to power supplies and the entire electrical system, to ensure not only emergency landings in the event of a failure, but also flight continuity (also known as "safe continuous flight and landing"). This is also known as "one-fail-safe," meaning "able to tolerate one failure."

[0004] These regulations raise many questions, especially regarding whether they can be done while maintaining reasonable costs. In fact, doubling everything would more than double the cost, as the aircraft would have to be upsized to accommodate the additional weight, and all of the flight capabilities of the heavier aircraft would have to be reassessed.

[0005] Therefore, other solutions must be found, most of which are based on two principles, sometimes used in combination: Very sophisticated energy management at the battery pack level, with dedicated modules that precisely manage the operating point of the battery pack. US Patent No. 9,586,690 describes a solution of this type. Using a power source to power the horizontal thrusters and only using batteries during takeoff and landing, this can improve the range and endurance of electric solutions. WO 2020 / 016510 and EP 3628593 describe solutions of this type. Summary of the Invention [Problem to be solved by the invention]

[0006] However, these solutions are not entirely satisfactory and in particular do not achieve a fail-safe solution. [Means for solving the problem]

[0007] The present invention aims to improve this situation. To this end, the present invention provides an aircraft with a hybrid power supply. The aircraft includes at least two horizontal drive thrusters, each driven by an electric motor, forming at least two horizontal drive units. The aircraft also includes at least four pairs of vertical take-off and landing rotors, each driven by an electric motor, and at least four power supplies connected to the corresponding electric motors of one pair of vertical take-off and landing rotors. Each pair of rotors forms a vertical drive unit together with its corresponding electric motor and power supply. Each vertical drive unit includes a power supply bus, and the output of the power supply bus is connectable to a single horizontal drive unit. The number of vertical drive units is such that each horizontal drive unit is connectable to at least two vertical drive units via a corresponding switch located at the input of the horizontal drive unit. The aircraft also includes at least two power generation sources. Each of the at least two power generation sources is connected to a respective power supply bus via the input of the corresponding vertical drive unit, on the one hand, and to each horizontal drive unit via the output of each vertical drive unit, on the other hand. The aircraft also includes at least one power supply control arranged to send power commands to the power generation source according to power requirements of the vertical drive unit and / or the horizontal drive unit. The power source supplies electricity according to a difference between the power requirements of the vertical drive unit and / or the horizontal drive unit and power output by the power generation source based on the power commands. The power generation source is adapted to recharge the power source such that the power source is passively controlled.

[0008] This aircraft is particularly advantageous because its architecture allows for redundancy to be created, ensuring one fail-safe while optimizing the size of the elements. In this way, the aircraft according to the invention minimizes the additional costs associated with realizing one fail-safe and achieves a true hybrid architecture with truly complementary power sources at each stage of flight.

[0009] According to various embodiments of the invention, the invention may have one or more of the following features. The vertical drive unit is powered by a battery. The power generation source includes a turbine generator and an AC / DC converter. The power generation source includes a hydrogen fuel cell and a DC / DC converter. The aircraft further comprises electrical contactors that connect the horizontal drive unit, the vertical drive unit, and the power generation source, respectively, to the remainder of the aircraft's electrical circuitry. Each element of the horizontal drive unit, vertical drive unit and power generation source is connected via an electrical contactor to the other element of the horizontal drive unit, vertical drive unit or power generation source to which it belongs. The aircraft further comprises diodes at the input and / or output of each of the horizontal drive unit, the vertical drive unit and the power generation source, connecting them to the electrical bus connecting them. The diode is placed upstream of the electrical contactor relative to the power generation source. [Brief explanation of the drawings]

[0010] Other characteristics and advantages of the invention will become more apparent from the following description, given for illustrative and non-limiting purposes, with reference to the accompanying drawings, in which: [Figure 1] 1 is a schematic diagram of an aircraft electrical architecture in accordance with the present invention; [Figure 2] FIG. 1 illustrates an example flight sequence showing which elements are active at which stage and the associated charge levels. [Figure 3] FIG. 10 is a diagram illustrating an example of a configuration in which a vertical drive unit is lost during takeoff and landing. [Figure 4] FIG. 10 is a diagram illustrating an example of a configuration in the event that a power generation source is lost. [Figure 5] FIG. 10 is a diagram showing a configuration example in which the horizontal drive unit is lost. [Figure 6] FIG. 1 illustrates an example of a power control management algorithm implemented on an aircraft. DETAILED DESCRIPTION OF THE INVENTION

[0011] The accompanying drawings and the following description substantially contain specific elements of the present invention, and therefore they not only serve to better understand the present invention but also contribute accordingly to its definition.

[0012] This specification may contain material that is subject to copyright protection. The Owner has no objection to the copying by anyone of the exact copy of this application or its description as published. The Owner otherwise reserves all rights thereto.

[0013] As shown in FIG. 1, an aircraft 2 according to the present invention comprises a control unit 4, two horizontal drive units 6 and 8, four vertical drive units 10, 12, 14 and 16, and two power generation sources 18 and 20.

[0014] In the embodiment described herein, the horizontal drive unit 6 (and 8) comprises a DC / AC converter 22 (and 32), an electric motor 24 (and 34), and a thruster 26 (and 36), e.g., having a propeller. The thruster 26 (and 36) is arranged to enable the aircraft to move forward in a substantially horizontal direction. In the embodiment described herein, the thruster 26 (and 36) consumes 80 kW of power during flight.

[0015] The horizontal drive unit 6 (and 8) is connected at its input to a switch 28 (and 38), which allows this input to be connected to the output of the vertical drive unit 10 (and 14) or 12 (and 16), as will be described below.

[0016] Vertical drive unit 10 (and 12, 14, 16) includes rotor 42 (and 46, 72, 76) driven by motor 52 (and 56, 82, 86) and rotor 44 (and 48, 74, 78) driven by motor 54 (and 58, 84, 88). Motors 52 and 54 are driven by DC / AC converters 62 and 64 (and 66 and 68, 92 and 94, 96 and 98), respectively. DC / AC converters 62 and 64 (and 66 and 68, 92 and 94, 96 and 98) are connected to an electrical bus of vertical drive unit 10 (and 12, 14, 16), which is connected to battery 50 (and 60, 80, 90), an input connected to the power distribution bus of power generation source 18, and an input connected to the power distribution bus of power generation source 20. The electric bus of each vertical drive unit 10 and 12 (and 14 and 16) is connected to a respective output of vertical drive unit 10 and 12 (and 14 and 16), which is connected to switch 28 (and 38). As will be described below, batteries 50, 60, 80 and 90 output a total of 600 kW at 100% capacity.

[0017] In the example described herein, each of the power generation sources 18 (and 20) includes a turbine generator 100 (and 102) and an AC / DC converter 104 (and 106). In the example described herein, each of the turbine generators can output 40 kW at 100% capacity. Alternatively, the power generation sources can be other DC or AC power generation sources followed by AC / DC or DC / DC converters. For example, these power generation sources can be based on turbogenerators powered by conventional fuels, biofuels, or synthetic fuels. Further alternatively, hydrogen-based energy sources, such as fuel cells, can be used.

[0018] As will be explained below, the control unit 4 is a low voltage device arranged to control the power generating sources 18 and 20 on the one hand and the switches 28 and 38 on the other hand, as well as various protective elements not shown in Figure 1, which will be explained below with reference to Figures 2 to 6.

[0019] Referring to Figure 1, we can see that all motors and electrical elements are duplicated. This ensures a one-fail-safe system, as will be explained later. In fact, there are two horizontal drive units, four vertical drive units connected to the horizontal drive units in two subgroups, and two power generation sources.

[0020] However, beyond this conventional duplication, the advantages of the present invention can be achieved by having a separate electrical bus for each vertical drive unit and a separate power distribution bus for each power generation source.

[0021] In fact, as we will see, the aircraft's special architecture, shown in Figure 1, allows for a true hybrid power supply, unlike existing solutions for parallel power supply arrangements. This allows both batteries and power generation sources to be used in conjunction, depending on the power requirements. Moreover, this architecture allows the battery to be treated simply as an "energy buffer." The battery is treated as a completely passive entity, requiring no software or hardware intelligence beyond the basic intelligence required for the battery system's operation by the Battery Management System (BMS), for example, to activate protection functions and report status. This completely contradicts all existing solutions in which the control role is played by an element specifically dedicated to optimizing battery operation, or an element dedicated to compensating for the battery's weaknesses, which operates exclusively alternately, i.e., the battery and this element never operate simultaneously.

[0022] FIG. 2 illustrates the energy consumption cycle during flight for the aircraft shown in FIG. 1. As shown, flight begins with the initial maneuver 200, when 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 operate for stability reasons, but consume only a small amount of energy. These are supplied with up to 600 kW from the battery and up to 80 kW from the power generation source. This reduces the battery capacity from 75% to 90% at the start of flight to 55% to 70%. At the end of this phase, the aircraft is approximately 50 feet, or 15 meters, above the takeoff point.

[0023] Then, in operation 210, the aircraft gradually transitions from vertical to horizontal flight between 50 and 150 feet, then climbs in a manner similar to a conventional aircraft. During this phase, the rotors are gradually shut down, and power consumption drops from 680 kW to 80 kW as the aircraft reaches horizontal cruise. The batteries and power generation sources continue to operate at full capacity. The batteries continue to deplete, dropping to 10% to 30% as the aircraft reaches horizontal cruise.

[0024] Horizontal flight is performed in operation 220, where the battery is not used. The power generation source continues to operate at full capacity, and the generated 80 kW is distributed among the horizontal drive units controlled by the control unit, and the power not consumed by the horizontal drive units is used to recharge the batteries. This phase allows the batteries to be recharged to about 50% at altitudes above 1000 feet (about 300 m).

[0025] Then, in operation 230, a ramp down is performed and the battery is recharged using 100% of the power generation source, which recharges the battery to 100%. During this phase, the motor elements do not consume energy.

[0026] Thereafter, a transition from horizontal flight to vertical flight is performed in operation 240. Here, power consumption gradually transitions to approximately 340 kW, and the battery charge gradually transitions from 100% to 85%-95%.

[0027] Finally, in operation 250, a vertical landing is performed using only the rotors as in takeoff, but also taking advantage of gravity, thereby continuing to discharge the battery to between 75% and 90%, as it was at the start of operation 200.

[0028] This improves the aircraft's usability by eliminating the need for recharging on the ground between two flights. Also, the power generation source is always fully operational (with exceptions as shown in Figure 6), and if the power generation source cannot provide enough power, the battery fills in. Similarly, the battery is recharged as quickly as possible to ensure sufficient electrical energy for landing.

[0029] Alternatively, the aircraft's batteries may be recharged on the ground between two flights sufficiently separated in time, in which case the control unit 4 may implement more sophisticated trade-offs between the power supply at different stages, such as changing the energy supply allocation to raise the operating point of the turbine generators to increase range, or to limit noise pollution or pollutant emissions.

[0030] The control of the battery is managed passively by the architecture of the present invention. If the rotor or horizontal drive unit consumes less than 80 kW, the battery is not used, of course, and the excess current can also be used to recharge the battery (e.g., as in operation 220 or 230). If more power is required, the battery is used, of course.

[0031] The architecture of the present invention also significantly reduces the need for larger batteries. In fact, it is not a coincidence that the battery has 10% to 30% charge remaining at the end of operation 220. This allows one fail-safe to be maintained even in the event of battery loss due to a malfunction.

[0032] Figure 3 shows an example of what would happen if the battery or other electrical components of one vertical drive unit failed. For simplicity, only the power generation source 18 and the energy output from the vertical drive unit 12 are shown, but the other components would operate similarly.

[0033] In the example shown, the first vertical drive unit 10 is switched off due to an electrical failure in one of the motors 52 and 54. It is disabled during one of operations 200, 210, 240 and 250.

[0034] First, note that each electrical element is protected by a controllable contactor that isolates it from the rest of the circuit. In this embodiment, two contactors (not shown) at the inputs of the vertical drive units 10 (and 12, 14, and 16) are connected to the power distribution bus specific to each power generation source, isolating the vertical drive units 10 (and 12, 14, and 16) to prevent electrical problems from propagating from the outside toward the vertical drive units 10 (and 12, 14, and 16). The same applies to the horizontal drive units with switches 28 and 38 and the power generation sources with switches (not shown). Each element in these electrical subsets is also connected to the rest of its electrical subset via a switch (not shown). This allows it to be isolated from the rest of the vertical drive units 10, even if the battery 50, for example, fails, without immediately being disconnected from the rest of the vertical drive units 10.

[0035] In the embodiment described herein, the contactors are duplicated by the presence of diodes (not shown), which can passively isolate the vertical drive unit in the event of an electrical problem, especially a short circuit, and prevent the propagation of the electrical problem from the vertical drive unit 10 (as well as 12, 14, 16) to the outside.

[0036] Additionally, batteries 50, 60, 80, and 90 are enlarged. Control unit 4 then causes switch 28 to connect to the output of vertical drive unit 12. This allows the batteries to be used up to 100 kW. Furthermore, the power generating source can be used for several minutes at 110% or 120%, exceeding its normal operating point. This, along with the overuse of the batteries, can compensate for the 80 kW loss caused by disconnecting vertical drive unit 10. The path taken by the electricity output from power generating source 18 is indicated by the bold arrow in the diagram.

[0037] This allows the battery to be slightly larger and still guarantee and perform takeoff (and emergency landing) without risk.

[0038] Figure 4 shows another failure example when one of the power generation sources is lost. Again, only some of the power elements are shown, but the other elements operate similarly.

[0039] As shown in the figure, the input of the vertical drive unit connected to the power distribution bus of the power generation source 10 is isolated by a contactor opened by the control unit 4. Also, the remaining power generation source is used above its normal operating point, and the battery is used to maintain a power supply of approximately 70 kW, i.e., 90% of the power normally consumed by the horizontal drive unit.

[0040] The batteries can be recharged during the descent phase before landing, ensuring one fail-safe here too, and a 25% larger battery ensures enough energy to continue level flight at 90% of normal capacity.

[0041] Figure 5 shows yet another example of a failure, this time involving the loss of a horizontal drive unit. In this case, the remaining horizontal drive unit is used to its full capacity to maintain horizontal cruise with an altitude and speed profile that accounts for the loss. The battery may be used as a buffer in case of sudden over-consumption.

[0042] FIG. 6 shows an algorithm that may be implemented by the control unit 4 to manage power commands in various situations.

[0043] This involves a cycle starting with operation 600, which receives the operating points of the rotors and / or thrusters.

[0044] Then, in operation 610, the rotor and / or thrusters draw a current corresponding to that operating point. This operation is followed in operation 620 by a test to determine whether the turbogenerator is operating at 100%. If not, in operation 630, the control unit 4 pushes the turbogenerator to maximum, and the battery supplements the required current while the turbogenerator increases speed. If the turbogenerator is boosted to maximum, in operation 640, the control unit 4 determines whether its power generation is sufficient to draw the current of operation 610. If not, full load is maintained until the next operating point, and the battery is loaded. If so, in operation 650, the control unit 4 verifies whether the battery needs charging. If charging is required, full load is maintained, and excess power is used to charge the battery until the next operating point. Otherwise, in operation 660, the control unit 4 decreases the operating point of the turbogenerator to the next operating point.

[0045] The power consumption figures given above are purely indicative and are not limiting, as the electrical architecture must be adapted to the actual requirements associated with flying the aircraft.

[0046] From the above, we can see that the power source is a high-power, low-capacity type, and the power generation source is a high-capacity, low-power type, because the power source is used as an energy buffer, and the power generation source is sized according to the power consumption during horizontal flight and to be able to charge the energy buffer.

[0047] This duality is achieved by the architecture described above, where all electrical circuits are interconnected and protected independently, making it possible to create redundancy that guarantees one fail-safe while optimizing the size of the elements. In this way, the aircraft according to the invention minimizes the additional costs associated with realizing one fail-safe and achieves a true hybrid architecture in which the power sources at each stage of flight are truly complementary.

[0048] It should also be noted that the diagram represents an electrical schematic of the aircraft. Therefore, while the diagram shows rotors 42 and 44, 46 and 48, 52 and 54, and 56 and 58 side-by-side, this is not necessarily shown from a mechanical perspective. In fact, the rotors are assembled in pairs on the vertical drive units to prevent a failure from causing aircraft instability. Therefore, rotors on the same vertical drive unit are generally positioned symmetrically about the center of the aircraft.

[0049] Also, while the diagram shows an aircraft with two thrusters, eight rotors, and two turbine generators, these numbers may be different. In fact, there may be more than two thrusters, and there may be more than two vertical drive units per horizontal drive unit. Similarly, the control unit 4 may be doubled for increased flexibility.

[0050] Finally, the architecture's control is simplified because one vertical drive unit can only be connected to one horizontal drive unit. In fact, this principle allows for redundancy and control of the magnitude by the number of different vertical drive units connected to the same horizontal drive unit. Furthermore, resiliency is easily achieved by adjusting the horizontal drive unit's switches. This is much more efficient than a system in which one vertical drive unit is connected to multiple horizontal drive units, which can be a significant control problem in both steady-state and degraded modes.

[0051] Alternatively, as mentioned above, the aircraft can be charged on the ground so that it is 100% powered at takeoff, which allows for other flight strategies, extends the aircraft's range, and reduces noise and pollutant emissions during low-altitude takeoff and landing.

Claims

1. 1. An aircraft having a hybrid power source, at least two horizontal drive thrusters (26, 36) driven by electric motors (24, 34), respectively, forming at least two horizontal drive units (6, 8); at least four sets of vertical take-off and landing rotors (42, 44, 46, 48, 72, 74, 76, 78) each driven by an electric motor (52, 54, 56, 58, 82, 84, 86, 88); and at least four power sources (50, 60, 80, 90) respectively connected to the corresponding electric motors (52, 54, 56, 58, 82, 84, 86, 88) of a corresponding set of the vertical take-off and landing rotors (42, 44, 46, 48, 72, 74, 76, 78), wherein each set of the rotors (42, 44, 46, 48, 72, 74, 76, 78) is driven by a corresponding electric motor (52, 54, 56, 58, 82, 84, 86, 88). a vertical take-off and landing rotor and power supply, the vertical take-off and landing rotor and power supply forming vertical drive units (10, 12, 14, 16) together with the horizontal drive units (6, 8), each of the vertical drive units (10, 12, 14, 16) having a power supply bus, the output of which is connectable to one of the horizontal drive units (6, 8), and the number of vertical drive units (10, 12, 14, 16) being such that each of the horizontal drive units (6, 8) is connectable to at least two of the vertical drive units (10, 12, 14, 16) via corresponding switches (28, 38) arranged at the inputs of the horizontal drive units (6, 8); at least two power generating sources (18, 20) connected on the one hand to each of the power supply buses via inputs of the corresponding vertical drive units (10, 12, 14, 16) and on the other hand to each of the horizontal drive units (6, 8) via corresponding outputs of the vertical drive units (10, 12, 14, 16); at least one power supply control (4) arranged to send power commands to the power generation source (18, 20) according to power requirements of the vertical drive units (10, 12, 14, 16) and / or the horizontal drive units (6, 8), wherein the power source (50, 60, 80, 90) supplies power according to a 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 output by the power generation source (18, 20) based on the power commands, the power generation source (18, 20) being suitable for recharging the power source (50, 60, 80, 90) such that the power source (50, 60, 80, 90) is passively controlled; An aircraft equipped with:

2. 2. The aircraft of claim 1, wherein the power source (50, 60, 80, 90) of the vertical drive unit (10, 12, 14, 16) is a battery.

3. 3. The aircraft of claim 1, wherein the power generation source (18, 20) comprises a turbine generator (100, 102) and an AC / DC converter (104, 106).

4. 3. The aircraft of claim 1, wherein the power generation source (18, 20) comprises a hydrogen fuel cell (100, 102) and a DC / DC converter (104, 106).

5. 5. The aircraft of claim 1, further comprising electrical contactors connecting the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16), and the power generation sources (18, 20), respectively, to the remainder of the aircraft's electrical circuitry.

6. 6. An aircraft according to claim 5, wherein each of the horizontal drive units (6, 8), the vertical drive units (10, 12, 14, 16) and the elements of the power generation source (18, 20) is connected by an electrical contactor to the other element of the horizontal drive unit (6, 8), the vertical drive unit (10, 12, 14, 16) or the power generation source (18, 20) to which it belongs.

7. 7. The aircraft according to claim 1, further comprising diodes at the 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 sources (18, 20) connected to an electrical bus connecting them.

8. 7. An aircraft as claimed in claim 5 or 6 when combined with claim 7, wherein the diode is located upstream of the electrical contactor relative to the power generation source.

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