aircraft

Abstract

Aircraft with at least one main rotor (13), at least one drive propeller (15) and a drive system, wherein the drive system a first electric motor which drives the main rotor (13) and draws electrical energy from a first drive battery, b. a second electric motor which drives the propulsion propeller (15) and draws electrical energy from a second propulsion battery, and c. a motor-generator unit comprising an internal combustion engine and at least one generator for generating electrical energy, which generator is electrically connected to the first drive battery and / or the second drive battery, wherein the internal combustion engine is a two-cylinder reciprocating piston engine with two cylinder-piston units in a tandem arrangement, wherein each cylinder-piston unit has a crankshaft and the crankshafts of both cylinder-piston units are mechanically coupled to one another, wherein at least one crankshaft, in particular both crankshafts each, is mechanically connected to a generator.

Description

[0001] The invention relates to an aircraft.

[0002] Aircraft are known in practice that have a main rotor for generating lift and at least one drive propeller for propulsion. Such aircraft utilize the principle of autorotation for the main rotor, so that with sufficient propulsion provided by the drive propeller, the main rotor rotates automatically or without artificial power supply, thus keeping the aircraft in the air. Such aircraft enable airborne locomotion with relatively low energy consumption.

[0003] Such aircraft are known, for example, from US 2018 0 065 741 A1, DE 10 2019 117 553 A1, US 2018 0 334 251 A1, US 2018 0 208 305 A1, EP 3 385 160 A1 and WO 2019 096 358 A2.

[0004] Current developments regarding climate-friendly mobility in urban environments and the technical advancement of purely electric propulsion systems have also produced prototypes of small, purely electrically powered aircraft. Such aircraft, or multicopters, often have multiple main rotors mounted around a passenger compartment, which, through appropriate control, maintain the aircraft in a stable flight attitude. These aircraft are climate-friendly because they can be operated emission-free. However, the disadvantages of these solutions are their limited range and the high space requirements on the ground, as the battery technology available in the short and medium term has only low power densities.

[0005] Against this background, the invention is based on the object of specifying an aircraft that can be operated in a climate-friendly manner and yet has a long range, can take off and land vertically and is quiet.

[0006] According to the invention, this object is achieved by the subject matter of patent claim 1.

[0007] The invention is based on the idea of ​​providing an aircraft with at least one main rotor, at least one drive propeller, and a drive system, wherein the drive system comprises a first electric motor, a second electric motor, and a motor-generator unit. The first electric motor drives the main rotor and draws electrical energy from a first drive battery. The second electric motor drives the drive propeller and draws electrical energy from a second drive battery. The motor-generator unit comprises an internal combustion engine and at least one generator for generating electrical energy, which generator is electrically connected to the first drive battery and / or the second drive battery.

[0008] The aircraft according to the invention is based on the basic idea of ​​increasing the range of an electrically powered aircraft using a motor-generator unit. The motor-generator unit serves solely to generate electrical energy, so that the actual propulsion power for the aircraft is provided entirely by electricity. To ensure the most reliable operation possible, the invention provides for each electric motor to be assigned its own drive battery. Thus, the main rotor and the drive propeller have separate drive trains, so that the aircraft can continue to operate safely even if individual components of a drive train fail.

[0009] In order to further increase the reliability, it is preferably provided that the motor-generator unit has two generators, wherein a first generator is electrically connected to the first drive battery and a second generator is electrically connected to the second drive battery.

[0010] This creates two independent drive trains that are only connected by the motor-generator unit. However, the capacity of the drive batteries is designed to be sufficient to at least land the aircraft safely if the combustion engine of the motor-generator unit fails. However, if another component of the drive train fails, such as one of the generators, one of the electric motors, or one of the drive batteries, the power of the second drive train remains fully intact. To achieve a longer flight duration in the event of failure of one drive battery and / or one of the electric motors, it can also be designed for the generators of the motor-generator unit to be electrically connected to both drive batteries.If one drive battery fails, the other drive battery can be recharged with a higher electrical output, increasing the overall flight time for an emergency landing. It is also possible for the generators to be additionally connected, or connectable, to the electric motors of the drive propeller and / or the main rotor via a bypass circuit, so that if one drive battery fails, the electrical energy generated in the generators can be fed directly to the electric motors of the drive propeller and / or the main rotor.

[0011] It is also possible for the main rotor to have no connection to the combustion engine. For example, the main rotor can be equipped without any drive and used only in autorotation mode. However, the main rotor can also be driven purely electrically, with the electric motor associated with the main rotor being powered, for example, by a drive battery that is not charged by the motor-generator unit but exclusively by an external power source. Such an external power source could be, for example, an electric charging station (plug-in hybrid technology).

[0012] The propulsion system is preferably a serial hybrid propulsion system. In other words, the motor-generator unit is connected in series with the propulsion batteries and the electric motors. The motor-generator unit does not directly drive the aircraft, but merely provides energy to generate electrical power, which is then drawn upon by the electric motors via the propulsion batteries as a buffer storage and converted into kinetic energy for the aircraft. This has the advantage that the combustion engine can be operated in an optimal operating range with minimal load changes, preferably at a fuel-air ratio of λ = 1. This allows the combustion engine to operate particularly cleanly, eliminating the need for complex exhaust gas aftertreatment. This reduces the weight of the combustion engine and manufacturing costs, which can be crucial for the realization of such an aircraft.

[0013] Another advantage of the serial hybrid propulsion system is that the additional motor-generator unit allows the propulsion batteries to be smaller than required for purely electric-powered aircraft. Since the propulsion batteries account for a large portion of the aircraft's weight, this allows for weight reduction. This, in turn, reduces the power required for flight and significantly increases the range of such an aircraft.

[0014] For the implementation of a serial hybrid propulsion system, it is particularly preferred if the combustion engine of the motor-generator unit exclusively drives the at least one generator. If the aircraft is equipped with a motor-generator unit that has two generators, which is particularly preferred, the combustion engine can exclusively drive the two generators. In other words, the combustion engine exclusively drives generators or one generator, but does not directly cause the aircraft to move.

[0015] According to the invention, the internal combustion engine is a two-cylinder reciprocating piston engine with two cylinder-piston units arranged in tandem. Such a two-cylinder reciprocating piston engine with a tandem arrangement is particularly compact and can be operated with very low vibration. Furthermore, the weight of the two-cylinder reciprocating piston engine is particularly low and its efficiency is high.

[0016] According to the invention, the two cylinder-piston units each have a crankshaft, and the crankshafts of both cylinder-piston units are mechanically coupled to one another. For example, the crankshafts can have toothed spur gears that mesh with one another. In this respect, it is particularly preferred that the spur gears rotate in opposite directions. The cylinder-piston units can each have cylinder axes that are spaced from one another by a distance that is smaller than the distance between the center axes of the crankshafts. This inwardly engaged interleaving of the cylinders ensures particularly smooth running, especially when starting the internal combustion engine.

[0017] According to the invention, at least one of the crankshafts is mechanically connected to a generator. This applies to a motor-generator unit that has a single generator. In a motor-generator unit that has two generators, it is preferred that both crankshafts are each mechanically connected to a generator.

[0018] To make the internal combustion engine particularly compact, it can preferably be provided with an under-engine camshaft. Alternatively, two overhead camshafts can be provided, so that the valves of the individual cylinder-piston units can be controlled independently of one another.

[0019] In general, a preferred variant of the invention can provide for the combustion engine to have a control system configured so that one cylinder-piston unit can continue to operate in the event of a failure of the other cylinder-piston unit. For example, a first cylinder-piston unit can continue to operate if, for example, the valve control system of the second cylinder-piston unit is defective. This contributes significantly to the reliability of the aircraft.

[0020] To minimize noise and vibrations, the motor-generator unit can also be arranged in an encapsulated housing. The housing can be spring-mounted, in particular rubber-mounted, in the aircraft. The combustion engine, together with the at least one generator, in particular the two generators, forms the motor-generator unit. Each generator of the motor-generator unit preferably has an electrical output of at least 40 kW, preferably at least 50 kW, preferably at least 60 kW. The electric motor for the drive propeller can have an electrical power consumption of 50 kW. The electric motor for the main rotor can have an electrical power consumption of 20 kW.

[0021] In a further preferred embodiment of the invention, the propulsion system comprises at least one fuel tank, in particular for a methanol-based fuel. Operating the propulsion system with methanol as fuel is particularly advantageous because it allows the vehicle's overall emissions to be significantly reduced. Methanol is a very easy-to-produce and synthesize fuel based on hydrocarbons and / or alcohol. In particular, the CO2 footprint for operating the aircraft can be balanced if the methanol is produced using renewable energy. This enables particularly climate-friendly operation of the aircraft.

[0022] The main rotor of the aircraft is preferably adapted to operate according to the principle of autorotation. For this purpose, the drive system may, in particular, include a coupling arranged between the main rotor and the main rotor's electric motor in such a way that the main rotor can be mechanically decoupled from the electric motor.

[0023] In particular, the aforementioned method can be used to ensure that the aircraft initially gains altitude during the takeoff phase by means of the driven main rotor. As soon as sufficient propulsion has been achieved with at least one drive propeller, i.e., a sufficient cruising speed has been reached, the main rotor's electric motor can be decoupled and switched off, allowing the main rotor to enter autorotation due to the propulsion provided by the drive propeller. This allows the aircraft to be held stably in the air solely by the drive propeller and the corresponding adjustment of the propulsion speed. An additional power supply for the main rotor is not required, thus reducing the aircraft's energy requirements.

[0024] It is also possible for the main rotor to be driven by the electric motor only during takeoff until sufficient kinetic energy has been built up for a vertical takeoff. For example, the main rotor can be equipped with a collective blade pitch system. The rotor blades are set to a low, low-drag, non-lift-generating angle of attack using collective blade pitch. The electric motor then introduces kinetic rotational energy into the main rotor. As soon as a certain limiting speed is reached, the angle of attack of the rotor blades is increased using collective blade pitch to generate sufficient lift for the aircraft to take off (a so-called jump start). At the same time, the electric motor can be switched off or the main rotor can be decoupled from the electric motor to reduce the risk of unwanted and dangerous rotation of the flooding device around its vertical axis.

[0025] It is particularly preferred if the aircraft has two drive propellers whose rotational axes are arranged parallel to each other. The drive propellers are preferably designed to rotate in opposite directions, thus compensating for any torque exerted by the main rotor on the passenger compartment. This ensures stable flight. The drive system can be designed as a 48-volt system. It is also possible to design the drive system as a 400-volt or 800-volt system. 400-volt or 800-volt systems can additionally incorporate a rapid-charging function so that the drive batteries can be charged with a high direct current via a rapid-charging station. In all cases, it is preferred that the aircraft be supplied with external electrical power to quickly charge the drive batteries.

[0026] The use of two drive propellers is particularly advantageous for vertical takeoff and landing. The drive propellers are preferably aligned so that their rotation axes run parallel and laterally to the aircraft. This compensates for any torque generated by driving the main rotor. Consequently, the main rotor can continue to be driven during takeoff until a predetermined flight situation, such as a minimum flight altitude and / or minimum flight speed, is reached. Similarly, for a vertical landing, driving the main rotor makes the vertical landing more controllable. The two drive propellers also compensate for the rotor torque that would otherwise act on the passenger compartment and cause it to rotate.

[0027] The electric motor for the main rotor can have a power consumption of 20 kW. Since propulsion during cruise flight is provided by the drive propellers, which keep the main rotor, which is decoupled from the electric motor, in autorotation, a relatively low-power electric motor is sufficient for the takeoff phase.

[0028] To achieve a safe vertical landing, it is advantageous if the aircraft, according to a preferred embodiment, has at least two counter-rotating propulsion propellers for generating propulsion, which are arranged on opposite sides of a fuselage axis of the aircraft. A first propulsion propeller can be driven by the second electric motor, and a second propulsion propeller can be driven by a third electric motor.

[0029] The aircraft preferably has at least two drive batteries, each electrically connected to the first electric motor, the second electric motor, and the third electric motor. The use of two drive batteries is advantageous for achieving redundancy. This redundancy significantly increases safety, especially during landing.

[0030] Preferably, the first electric motor is adapted to temporarily drive the main rotor and has an electrical power of no more than 130 W / kg, in particular no more than 125 W / kg, in particular no more than 100 W / kg, in particular no more than 90 W / kg, based on the takeoff mass of the aircraft. The counter-rotating drive propellers compensate for the torque of the main rotor and thus ensure a stable flight attitude, especially during landing.

[0031] The main rotor is preferably only driven temporarily by the first electric motor. In other words, the first electric motor can be decoupled from the main rotor or the first electric motor can be switched off, allowing the main rotor to continue rotating passively. The main rotor is therefore not driven during cruise flight but uses the principle of autorotation to generate lift. However, for landing, the first electric motor can be used to briefly supply additional power to the main rotor to assist lift (assist power), sufficient to enable a safe, nearly vertical landing. Specifically, the additional power supply can be used to delay a drop in speed of the main rotor, thus extending the time window for a soft landing. A relatively low-power electric motor is sufficient for this. This, in turn, advantageously reduces the weight of the aircraft.Specifically, the power of the electric motor is preferably dimensioned such that hovering using the first electric motor and the main rotor is not possible, but the power of the first electric motor is sufficient to enable a safe, almost vertical landing.

[0032] Theoretically, it is possible to land an aircraft vertically without main rotor propulsion. During cruise flight, the main rotor receives sufficient power from the aircraft's forward thrust to maintain rotation. This creates lift during cruise flight. If thrust is then significantly reduced to initiate a vertical landing, the aircraft will enter a descent of approximately 7 m / s. The descent rate is therefore so high that it would result in a hard landing, which could result in damage to the aircraft or injury to the aircraft's occupants. Taking into account the so-called ground effect, which creates additional lift, a time window of only 0.56 seconds remains for a smooth vertical landing using only the power of the main rotor generated by autorotation. This applies, for example, to an aircraft with a main rotor moment of inertia of 250 kgm. 2, an initial speed of 480 rpm at the beginning of the landing maneuver and a final speed of 300 rpm at the end of the landing maneuver. The final speed results from the power loss in the main rotor when initiating the landing maneuver. If the previously described time window is exceeded, the descent rate increases to a value too high for a smooth landing. The available time window is therefore too short to actually land the aircraft safely.

[0033] Using the first electric motor to supply support power to the main rotor therefore extends this time. For example, in an aircraft with a takeoff weight of 600 kg, the first electric motor can supply 50 kW of electrical power to the main rotor, almost tripling the time available for landing. In the example described above (main rotor moment of inertia 250 kgm 2, initial speed 480 rpm and final speed 300 rpm), the time window available for a soft landing increases to 1.48 seconds. To ensure that sufficient power can be transferred to the main rotor for a nearly vertical landing, it is preferably provided that the first electric motor has an electrical power of at least 60 W / kg, in particular at least 70 W / kg, in particular at least 80 W / kg. The aforementioned values ​​apply in each case with respect to the take-off mass of the aircraft.

[0034] Even a time window of 0.5 seconds or 1 second is too short for a manual landing of the aircraft, although it is possible. Therefore, flight safety is enhanced if, as provided for in a preferred embodiment of the invention, a flight control system is provided. The flight control system is preferably connected to a flight altitude sensor, a rotor blade adjustment system of the main rotor, and the first, second, and third electric motors in such a way that, for landing the aircraft, a. if the altitude sensor detects that the aircraft falls below a predetermined first landing altitude, the main rotor is driven by the first electric motor to provide lift support, the second and third electric motors being controlled in such a way that a torque of the main rotor caused by the lift support in the main rotor is compensated, and b. if the altitude sensor detects that the aircraft falls below a predetermined second landing altitude, the rotor blade pitch is adjusted so that the aircraft's descent speed is reduced, in particular to a value of no more than 0.3 m / s.

[0035] The flight control system has the decisive advantage of enabling a soft landing with relatively low energy expenditure. By supplying assist power to the main rotor, the time window for a near-vertical landing is extended, allowing for better control and a smoother landing. Specifically, the assist power reduces the aircraft's rate of descent during the first landing phase. In the second landing phase, when the aircraft falls below the second landing altitude, the rotor blade adjustment further reduces the rate of descent, allowing the aircraft to touch down gently on a landing surface. The second landing altitude is lower than the first landing altitude.

[0036] In this respect, according to a subordinate aspect of the invention, a method for controlling a previously described aircraft is also disclosed and claimed, the method comprising the following steps: a. detecting a drop below a predetermined first landing altitude by means of an altitude sensor; b. driving the main rotor (13) by means of the first electric motor when the first landing altitude is determined, wherein the second and third electric motors are simultaneously controlled in such a way that an additional torque generated by the first electric motor in the main rotor (13) is compensated; c. detecting a drop below a predetermined second landing altitude by means of the altitude sensor; d. Adjusting the rotor blade pitch when the aircraft falls below the second landing altitude so that the rate of descent of the aircraft is reduced, in particular to a value of no more than 0.3 m / s.

[0037] The previously described method is particularly simple and utilizes a small landing area. This makes it possible to land the aircraft, for example, on the roof of a high-rise building, even though there is little available landing area. For the previously described landing, the aircraft preferably requires a landing area of ​​no more than 10 m. In the context of this application, a landing is also referred to as a vertical landing if a landing area of ​​less than 10 m is sufficient.

[0038] Crucial for the aircraft's particularly favorable landing in confined spaces is the provision of support power to the main rotor by the first electric motor. For landing, the main rotor receives additional power from the first electric motor. Through rotor blade adjustment, this kinetic rotational energy can then be converted into controlled lift, with the lift being controlled in such a way that the aircraft transitions into a gentle landing approach. This preferably occurs immediately before landing and by utilizing ground effect.

[0039] Furthermore, an aircraft with a flight control system is preferably described, which enables particularly simple and responsive flight control, in particular for takeoff and landing and for autopilot operation. The flight control system can have a flight altitude sensor.

[0040] Preferably, a flight control system is provided for one of the previously described electric aircraft. The flight control system comprises electrically driven flight control surfaces, an electrically driven main rotor, and two drive propellers, each driven by two independent electric motors. The electric motors of the flight control surfaces, the main rotor, and the drive propellers are connected to a control unit that has attitude sensors and is configured to control the individual electric motors such that the aircraft assumes a predetermined flight attitude. The current flight attitude determined by the attitude sensors is compared with a predetermined desired flight attitude.

[0041] Preferably, all components relevant for attitude control, in particular the tilt head control of the rotor, elevator, rudder, drive propeller, and main rotor, are electrically operated and controllable. This allows autopilot control to be carried out entirely electrically. For this purpose, the flight control system uses several sensors and, preferably, additional data. The additional data can include position data determined, for example, by GPS or GLONASS. Furthermore, environmental data, such as air pressure, altitude, and airspeed, can be incorporated into the flight control system. Data from Wi-Fi networks can also be used, particularly for positioning. Furthermore, environmental sensors, such as ultrasonic sensors or radar sensors, can detect other objects, and the flight control can be adjusted accordingly in the flight control system.For example, ultrasonic sensors can be used to measure the distance to the ground, allowing the aircraft to land safely and automatically. Radar sensors can ensure that collisions with other aircraft are avoided. Another data source can be air traffic data, which contains the position data or transponder signals of commercial aircraft. This can ensure that the aircraft autonomously flies well around the flight paths of larger aircraft, for example, to avoid turbulence caused by wake turbulence.

[0042] The flight control system can also be connected to an electrically operated clutch arranged between the main rotor and the main rotor's first electric motor. Alternatively, a freewheel can be provided between the first electric motor and the main rotor. This allows the main rotor to be decoupled from its driving electric motor upon reaching a certain flight speed. The aircraft thus automatically transitions from an operating state in which the main rotor actively provides lift to autorotation mode, in which the main rotor is kept rotating by the forward movement of the aircraft.

[0043] It is preferred if the control unit of the flight control system has a data interface to a navigation system and is configured such that, based on the data from the navigation system, an autonomous flight from a predetermined starting point to a predetermined destination can be carried out. The data from the navigation system can be, for example, GPS data or GLONASS data. Other data that include position determination is also possible. The navigation system can also have map data, wherein the map data preferably also includes elevation data and special landmarks relevant to flight operations. Such landmarks can be, for example, high-voltage power lines, commercial airports, tall buildings, or wind turbines.

[0044] The flight control system may also include sensors and / or communicate with sensors at a landing site to enable fully automatic landing of the aircraft.

[0045] Preferably, an aircraft is described that has a main rotor comprising a plurality of adjustable rotor blades. The main rotor may further comprise at least one electromagnetic actuator coupled to the adjustable rotor blades to set a predetermined pitch angle of the rotor blades.

[0046] The lift of the main rotor can be influenced by varying the rotor blade pitch angles. Therefore, it is preferable that the rotor blades be adjusted for takeoff to generate the highest possible lift, particularly maximum lift. During flight, especially at cruising speed, the rotor blade pitch angle can be changed to operate the main rotor in autorotation mode. The adjustable rotor blades thus enable particularly efficient flight operations.

[0047] In a preferred variant, the predetermined adjustment angle has three discrete adjustment angles. In particular, a controller can be provided that is coupled to the electromagnetic actuator such that the electromagnetic actuator can position the rotor blades at three discrete adjustment angles. In particular, it can be provided that a first adjustment angle is less than 0 degrees, a second adjustment angle is at most 2.5 degrees, and a third adjustment angle is more than 2.5 degrees.

[0048] The adjustment angles mentioned enable different flight functions.

[0049] The first pitch angle, which is less than 0 degrees or negative, can be used in situations where no lift is required. A pitch angle of less than 0 degrees ensures that the main rotor rotates without lift. The flow resistance is low, so this position is used in particular to start the main rotor with high energy efficiency. When the main rotor starts up, as much of the power of the electric motor that drives the main rotor as possible is converted into kinetic energy, which is stored in the rotational movement of the main rotor for a subsequent start, in particular a so-called jump start. By adjusting the rotor blades to a negative pitch angle, the electrical energy is used particularly efficiently to generate kinetic rotational energy in the main rotor, as the air resistance of the individual rotor blades is reduced.It can also be useful to generate a slight downforce so that the aircraft remains wind-stable when the main rotor starts running.

[0050] For the subsequent (jump) takeoff, once sufficient rotational energy has been generated in the main rotor, the rotor blades can be adjusted to a pitch angle of more than 2.5 degrees. The kinetic rotational energy stored in the main rotor is thus converted into lift, allowing the aircraft to take off essentially vertically. The main rotor's electric motor continues to operate to assist the takeoff process.

[0051] By applying forward thrust as soon as possible after the aircraft takes off, preferably generated by the at least one drive propeller, the aircraft begins to move forward. As soon as a predetermined minimum speed is reached, the main rotor can then be switched to autorotation mode. For autorotation mode, a second rotor blade pitch angle of no more than 2.5 degrees is preferably set.

[0052] Preferably, the adjustable rotor blades are spring-loaded such that, when the electromagnetic actuator is inactive, the adjustable rotor blades assume a predetermined adjustment angle and are locked. It is particularly preferred if the adjustable rotor blades automatically assume and maintain the second adjustment angle when the electromagnetic actuator is inactive.

[0053] In other words, the main rotor's autorotation mode can be set by de-energizing the electromagnetic actuator. The spring preload then ensures that the rotor blades automatically assume the second pitch angle and thus the position for autorotation mode. In this position, the rotor blades are preferably mechanically locked.

[0054] This sets the aircraft apart from previous aircraft that enable jump starts. In known aircraft that use jump starts, the main rotor is only powered as long as the aircraft is in contact with the ground. As soon as the aircraft leaves the ground, the main rotor drive is deactivated. With the invention, the main rotor drive can remain activated, which leads to improved takeoff behavior. The resulting torques on the passenger compartment, which are triggered by the main rotor, are compensated for by the drive propellers. This is achieved through asymmetric thrust from the drive propellers. Accordingly, the main rotor's electric motor can continue to operate even when the aircraft is landing. This enables particularly good control of the landing, and in particular, a nearly vertical landing can be achieved.The space required for takeoff and landing is therefore particularly small, which is of great importance for use in an urban environment.

[0055] The aircraft preferably has two counter-rotating propellers that compensate for the torque of the main rotor.

[0056] Generally, all blades of the main rotor are also designed to be simultaneously adjustable. The main rotor may have at least two or more blades.

[0057] In general, it is possible for all aircraft described in this application to use two counter-rotating main rotors instead of or in addition to the two drive propellers. The counter-rotating main rotors already sufficiently balance the rotor torque acting on the passenger compartment.

[0058] Preferably, an aircraft is described with two drive propellers, each of which is electrically driven and encapsulated in a soundproof ring.

[0059] The noise protection ring can be formed by one or more wings of the aircraft.

[0060] When using aircraft, particularly in urban environments, it is necessary to generate as little noise as possible. This applies in particular to the takeoff and landing phases of the aircraft. Therefore, the invention provides for the aircraft to have drive batteries that can store sufficient electrical energy to complete a takeoff and landing. In other words, the aircraft should be powered purely by battery-electric power during the takeoff and landing phases. This already significantly reduces noise emissions. However, the drive propellers generate noise due to vortex effects. To reduce this, the invention provides for the drive propellers to be encapsulated in a soundproofing ring. This significantly reduces noise emissions, at least to the sides of the aircraft.The drive propellers, each encapsulated in a soundproofing ring, can be used in all aircraft described in this application.

[0061] For all embodiments of aircraft in this application, the aircraft is preferably designed as a sports aircraft.

[0062] The invention will be explained in more detail below using an exemplary embodiment with reference to the attached schematic drawings. Fig. 1 a front view of an aircraft according to the invention according to a preferred embodiment; Fig. 2 a side view of the aircraft according to Fig. 1; Fig. 3 a plan view of the aircraft according to Fig. 1; Fig. 4 a plan view of an aircraft according to the invention according to a further preferred embodiment; Fig. 5 a side view of the aircraft according to Fig. 4; Fig. 6 a cross-sectional view of the aircraft according to Fig. 4; Fig. 7 a cross-sectional view of a power unit for an aircraft according to the invention according to a preferred embodiment and Fig. 8 a perspective view of a drive battery for an aircraft according to the invention according to a preferred embodiment.

[0063] The aircraft according to the invention is preferably designed as a small aircraft for a maximum of four, in particular a maximum of three, and in particular a maximum of two people. This type of aircraft can essentially be a combination helicopter. The aircraft can be of lightweight construction. In particular, the aircraft can be designed and approved as a sports aircraft.

[0064] Overall, the aircraft comprises a passenger cell 10 connected to wings 11. Front wings 11a and rear wings 11b are provided. The rear wings 11b are mounted higher on the passenger cell 10 than the front wings 11a. The rear wings 11b can also have downward-facing winglets 11e. Furthermore, the rear wings 11b can have a greater span than the front wings 11a. The passenger cell 10 includes a cockpit 12.

[0065] A main rotor 13 is also mounted on the passenger compartment 10. The main rotor 13 has two rotor blades 14. The main rotor 13 is driven by a first electric motor. Preferably, a coupling is provided between the first electric motor and the main rotor 13, so that the main rotor 13 can be completely decoupled from the first electric motor. Alternatively, the first electric motor can be designed as a brushless and / or separately excited motor, so that complete mechanical decoupling of the main rotor 13 can be achieved by de-energizing the first electric motor.

[0066] Two drive propellers 15 are also mounted on the rear wing 11. The drive propellers 15 are aligned with their rotation axes parallel to each other. The drive propellers 15 are attached to the rear of the rear wing 11. The drive propellers 15 are preferably each driven by an electric motor, which can be integrated into the drive propeller 15. Specifically, a first drive propeller 15' can be driven by a second electric motor, and a second drive propeller 15" can be driven by a third electric motor.

[0067] In order to reduce the noise emissions of the drive propellers 15, the drive propellers 15 each have soundproofing rings 16 in which the drive propellers 15 are arranged. To illustrate the arrangement of the drive propeller 15 within the soundproofing ring 16, Fig. 3 one of the two soundproofing rings 16 is shown in a partial section.

[0068] For flight control, the aircraft also has rudders 17a, each arranged behind the drive propellers 15. Not shown in the drawings are elevators, which are preferably located on at least one of the wings 11, preferably on both wings 11. Additionally, ailerons may also be provided, which are usually arranged laterally near the winglets 11e on the rear wing 11b.

[0069] In the aircraft shown here, the elevator and aileron functions can be achieved by tilting the rotor head. The arrangement of the drive propellers 15 and rudder 17a represents a compromise. On the one hand, the lateral, parallel arrangement of the drive propellers 15 compensates for the torque generated when driving the main rotor 13. On the other hand, the two drive propellers 15 may only be positioned far enough from the center of the aircraft so that, in the event of a failure of one of the two drive propellers 15, the now asymmetrical propulsion can be compensated by the two rudders 17a, and the aircraft remains fully controllable.

[0070] In the Fig. Not shown in Figures 1 to 3 is an additional motor-generator unit with an internal combustion engine, preferably designed as a two-cylinder reciprocating piston engine. The two-cylinder reciprocating piston engine comprises two cylinder-piston units operated in tandem and having counter-rotating, coupled crankshafts. The internal combustion engine is preferably operated within an optimal speed range, preferably at a particularly constant speed, so that pollutant emissions are low. Pollutant aftertreatment using a particulate filter or catalytic converter, etc., is not required.

[0071] The combustion engine is also coupled to two generators, each mechanically connected to a crankshaft. The generators provide electrical energy, which is transferred to the propulsion batteries integrated into the aircraft. The propulsion batteries are dimensioned to enable purely electric takeoff and landing, with the combustion engine shutting down during these phases of flight.

[0072] During cruise flight, the motor-generator unit is preferably operated and, via the generators, provides the electrical energy required for flight operation to utilize the electric motors. A 48-volt, 400-volt, or 800-volt system is preferably integrated as the electrical system. The motor-generator unit can be mounted, in particular, in the rear of the aircraft. The drive batteries are preferably mounted centrally in the area below the main rotor 13 to ensure good aircraft balance. A fuel tank for the fuel required for the combustion engine can be integrated into one or more wings 11. The fuel tank is preferably suitable for methanol as a fuel.

[0073] The motor-generator unit can also be designed to be so low-noise and low-vibration, or encapsulated against noise and vibration, that passengers cannot tell whether the combustion engine is activated or deactivated during flight. The design of the two-cylinder reciprocating piston engine is particularly well-suited for such low-noise and low-vibration operation. In particular, the two-cylinder reciprocating piston engine can be designed to be particularly low-vibration by interlacing the cylinder-piston units so that the cylinder axes are closer to each other than the crankshaft axes. This minimizes starting mass moments, particularly during combustion engine start-up, preventing the otherwise familiar starting vibrations. The combustion engine is preferably controlled so that the capacity of the electric drive batteries is maintained at approximately 80% during cruise flight.This protects the drive batteries and ensures that sufficient electrical energy is available for a purely electric landing.

[0074] In the Fig. 4 to 6, an aircraft according to a further preferred embodiment is shown. This type of aircraft is essentially a combination helicopter and / or gyrocopter. The aircraft also comprises a passenger cabin 10 with a cockpit 12. Furthermore, wings 11 are provided, which are fixedly arranged on the passenger cabin 10. Front wings 11a form in the aircraft according to Fig. 4 to 6 a horizontal stabilizer. Rear wings 11b essentially form the main wings. As shown in Fig. As can be seen in Figure 6, two propulsion propellers 15 are also provided, mounted on nacelles 15a. The nacelles 15a are firmly connected to the passenger compartment.

[0075] The aircraft according to Fig. 4 to 6 also includes a landing gear 18 with a rotating nose wheel 18a and two tail wheels 18b. The nose wheel 18a and the tail wheels 18b can be retractable.

[0076] Just as with aircraft according to Fig. 1 and Fig. 2, the aircraft shall, in accordance with Fig. 4 to 6 additionally have a main rotor 13 comprising two rotor blades 14. The rotor blades 14 are adjustable so that their angle of attack can be adapted. The adjustment of the rotor blades 14 is preferably carried out via a corresponding mechanism, which is referred to as rotor blade adjustment. The main rotor 13 is firmly connected to the passenger compartment 10 via a rotor boom 13a. The main rotor 13 can be tilted against and transversely to the direction of flight. This arrangement is called tilt-head control. The tilting of the main rotor 13 is achieved via the rotor boom 13a. Specifically, control rods can be guided in the rotor boom 13a, which enable tilting of the main rotor 13.

[0077] The main rotor 13 and the drive propellers 15 are each driven by electric motors. Preferably, a coupling is provided between the first electric motor of the main rotor 13 and the main rotor 13 itself, so that the main rotor 13 can be completely decoupled from the first electric motor. Alternatively, the first electric motor can be designed as a brushless and / or separately excited motor, so that complete mechanical decoupling of the main rotor 13 can be achieved by de-energizing the first electric motor.

[0078] During operation, the main rotor can be driven by an electric motor, particularly to enable a fast, nearly vertical takeoff (so-called jump start). During actual flight operation, the aircraft is preferably driven exclusively by the drive propellers 15. For this purpose, each drive propeller 15 is connected to another electric motor. Specifically, a second electric motor is assigned to a first drive propeller 15', and a third electric motor is assigned to a second drive propeller 15". The propulsion automatically sets the main rotor 13 in rotation (autorotation). The first electric motor of the main rotor 13 can thus be switched off. This enables particularly efficient flight operation.

[0079] In the cross-sectional view according to Fig. The structure of the rear wing 11b is clearly visible in Figure 5. The rear wing 11b essentially forms a frame that surrounds the drive propellers 15. Specifically, the rear wing 11b has an upper wing 11c and a lower wing 11d, with the upper wing extending downwards at each of its ends and connected to the lower wing 11d. The tail wheels 18b are directly connected to the downwardly directed side sections of the upper wing 11c or are rotatably mounted in these side sections, thus providing optimal aerodynamic coverage.

[0080] The frame-like enclosure of the drive propellers 15 by the rear wing 11b significantly reduces noise emissions. Essentially, the drive propellers 15 are encapsulated in this way, deflecting the sound generated by the rotation of the drive propellers 15, thus reducing noise emissions, particularly near the ground. The rear wing 11b thus forms a soundproof enclosure for the drive propellers 15.

[0081] The laterally downward-directed sections of the upper wing 11c can form a vertical stabilizer 17 and preferably each comprise a rudder 17a.

[0082] In Fig. Figure 7 shows a power generator 20 for an aircraft according to the invention. The power generator 20 comprises a two-cylinder reciprocating piston engine 21 with a first cylinder-piston unit 22 and a second cylinder-piston unit 23. Each of the cylinder-piston units 22, 23 comprises a piston 24 guided in a cylinder 25. The piston 24 is coupled to a connecting rod 26, which connects the piston 24 to a crankshaft 27. The crankshafts 27 are aligned parallel to one another and each carry spur gears 27a with external teeth. The externally toothed spur gears 27a mesh with one another, so that the spur gears 27a rotate in opposite directions.

[0083] Each spur gear 27a is coupled to a generator 30 via a toothed belt 28. A total of two generators 30 are provided. The generators 30 also include counterweights 30a, which balance the inertial forces and inertial moments.

[0084] The power unit further comprises a cam belt 31 connecting one of the crankshafts 27 to camshafts 32. Each cylinder-piston unit 22, 23 is assigned a camshaft 32. The camshafts 32 each act on valves 33, with each cylinder-piston unit preferably having four valves 33.

[0085] Finally, an oil pan 34 is provided, in which an oil pump 35 is arranged. The oil pump 35 is driven by an oil pump belt 36, which connects the oil pump 35 to one of the crankshafts 27. Preferably, the oil pump 35 is connected to a different crankshaft 27 than the camshafts 32. Furthermore, an oil filter 37 is arranged on the oil pan 34.

[0086] As in Fig. As can be clearly seen in Figure 7, the power generator 20 has a particularly compact design. It requires relatively few parts and is therefore easy to maintain and lightweight. The two-cylinder reciprocating piston engine is particularly quiet and low in vibration. In addition, the two-cylinder reciprocating piston engine can be encapsulated in a housing, whereby the housing can further contribute to the low noise and vibration levels. Preferably, the two-cylinder reciprocating piston engine is designed in such a way that passengers cannot detect whether the power generator 20 is activated or deactivated during flight.

[0087] In Fig. 7 also shows that the cylinders 25 of the cylinder-piston units 21, 22 are arranged offset from one another. In particular, the center axes of the cylinders 25 are spaced apart by a smaller distance than the center axes of the crankshafts 27. As a result, the connecting rods 26 are at the top dead center of the pistons 24, as shown in Fig. 7, are slightly inclined relative to each other. This significantly reduces vibrations during engine start-up. This reduces starting mass moments, particularly during the start-up of generator set 20, thus preventing the otherwise familiar starting vibrations.

[0088] During cruise flight, the power generator 20 is preferably operated and, via the generators 30, provides the electrical energy required for flight operation to utilize the electric motors. The electrical system preferably includes a 48-volt, 400-volt, or 800-volt system. The propulsion batteries are preferably mounted in the aircraft so that the aircraft's center of gravity is well balanced. Fuel tanks for the fuel required to operate the power generator 20 can be provided in the wings 11. The power generator 20 is preferably powered by methanol.

[0089] However, the use of fossil fuels should only occur during cruise flight. This is ensured by appropriate control. In particular, the drive batteries are dimensioned to enable purely electric takeoff and landing. During these flight phases, the power generator is therefore switched off. The power generator, however, is preferably activated when the battery charge level drops below a predetermined threshold and / or a minimum flight altitude is reached. It is particularly advantageous if the power generator 20 is controlled such that the capacity of the electric drive batteries is maintained at approximately 80% during cruise flight. This protects the drive batteries and simultaneously ensures that sufficient electrical energy is available for a purely electric landing.

[0090] Fig. Figure 8 shows a drive battery of an aircraft according to the invention according to a preferred embodiment. The drive battery comprises one, preferably a single, cell block 120 formed from a plurality of battery cells 121. The battery cells 121 are preferably arranged adjacent to one another in staggered rows with a low packing density. Lithium-ion round cells, preferably of the 18650 / 2170 type, are used as battery cells 121. The cell block 120 can have an electrical voltage of 48 volts or 60 volts. The electrical energy density is preferably between 2.1 kWh and 3.3 kWh. The battery system preferably has a footprint of 200 x 200 mm.

[0091] The battery cells 121 are electrically and mechanically connected by contact plates 122. The contact plates 122 extend over the battery terminals and are welded to them, preferably by laser welding. A contact plate 122 connects two rows of battery cells 121.

[0092] An electronics board 123 is arranged laterally along the cell block 120. The electronics board 123 includes a battery monitoring system and a plurality of clamp contacts that are electrically connected to the contact plates 122. The clamp contacts are arranged so that each row of battery cells 121 can be monitored individually.

[0093] The cell block 120 is encased in an insulating sheath 124. The insulating sheath 124 is made of an electrically insulating and thermally conductive material. In particular, the insulating sheath 124 is formed by a flexible film that can fit tightly against the cell block 120. The insulating sheath 124 encases the cell block 120 and connection modules 125 arranged on the end faces of the cell block 120. The connection modules 125 carry the essential electrical and, if applicable, pneumatic or hydraulic connections for connecting the battery system to external components.

[0094] Access openings 125a for connecting a vacuum pump 133 are arranged in a front-end connection module of cell block 120. The access openings 125a open within the insulating casing. The vacuum pump 133 is preferably electrically connectable to the cell block 120 and, in particular, operable with the rated voltage (48V, 400V, or 800V) of the cell block 120. The vacuum pump 133 is connected to the electronics board 123, in particular to the battery monitoring or management system (BMS).

[0095] The housing 110 of the battery system is formed from an aluminum sheet with a wall thickness of preferably 1 mm. The housing 110 has two housing end walls 112 that cover the connection modules. The side surfaces of the housing 110 are formed by housing side walls 111, which are welded to the housing end walls 112 in a gas-tight manner. The housing side walls 111 enclose the cell block 120. Preferably, the housing side walls 111 lie directly against the insulating casing 124 of the cell block 120.

[0096] As in Fig. As can be seen in Figure 8, an upper housing side wall 111 has two projections 111a with mounting holes 111b. This allows the battery system to be easily installed in the aircraft.

[0097] The housing side wall 111, in particular the upper and / or lower housing side wall 111, can be equipped with a cooling element 140. The cooling element 140 can be formed by a corrugated aluminum sheet structure. The cooling element 140 is preferably attached to the housing side wall 111 with good thermal conductivity.

[0098] The housing 110 has at least one connection 130 that extends into the housing 110 and is connectable or connected to the vacuum pump 133 or vacuum pump. The connection 130 is arranged in particular in the housing end wall 112. Through the connection to the vacuum pump, a negative pressure can be established within the housing 110. The negative pressure causes the flexible insulating sheath 124 to deform and fit tightly against the cell block 120. In the embodiment shown here, three connections 130 are provided in the housing end wall 112.

[0099] In general, at least one connection 130 may include a check valve, so that the negative pressure remains within the housing 110 after the vacuum pump is switched off. For safety reasons, it is also expedient for the housing 110 to be further equipped with a pressure relief valve 132. Such a pressure relief valve 132 opens to the outside when a predetermined pressure within the housing 10 is exceeded.

[0100] The housing side wall 111, in particular the upper and / or lower housing side wall 111, exhibits a certain degree of flexibility due to the material selection (aluminum) and the thin wall thickness (1 mm) and is also deformed by the negative pressure. As a result, the housing side wall 111 fits tightly against the insulating sheath 124 that tightly surrounds the cell block 120. The housing 110 is thus clamped to the cell block 120.

[0101] The housing end wall 112 further supports an electrical connector 131 for electrically connecting the battery system to external components. The electrical connector 131 can comprise a data line, in particular for connecting to a master / slave bus system. The electrical connector 131 is preferably arranged off-center and offset in height to ensure correct electrical connection to external components during installation. The battery system preferably comprises two differently polarized electrical connectors 131 arranged in opposite housing end walls 112. For example, the positive terminal connector can be arranged on a front housing end wall 112 and the negative terminal connector can be arranged on a rear housing end wall 112. This creates a large protective distance between the electrical connectors 131, which is expedient due to the high current intensities. Furthermore, this increases the protection against reverse polarity.

[0102] The vacuum pump, which is designed to maintain the negative pressure within the housing 110, can be connected at least indirectly to the electrical connection 131. The vacuum pump is thus supplied with the necessary operating voltage by the battery system itself, so that the entire system is self-sufficient.

[0103] Furthermore, it is provided that the vacuum pump is activated in a self-regulating manner. For this purpose, the battery system has a pressure sensor arranged within the housing 110. The pressure sensor, which is connected to a corresponding control unit, continuously monitors the vacuum within the housing 110. As soon as the vacuum leaves a predetermined target value range or falls below a predetermined target value, the control unit sends a control signal to the vacuum pump, activating the vacuum pump. If the vacuum subsequently reaches the predetermined target value again, the control unit sends another control signal to stop the vacuum pump.

[0104] Safety monitoring can also be implemented using the pressure sensor. If, for example, the control unit detects that the negative pressure cannot be maintained for a sufficient length of time or that the negative pressure leaves the predetermined target value range too frequently or too quickly, it can be assumed that the housing 110 is leaking. The control unit then sends a control command that initiates, for example, the output of an alarm signal and / or the shutdown of the battery system. Furthermore, the control signal can activate a second battery system, which can be provided as a backup module in a vehicle, in particular an aircraft. The second battery system is preferably integrated into the master / slave bus system.

[0105] The control unit can also output corresponding control signals, particularly with regard to safety-relevant events, based on data from the battery monitoring system. For example, the output of an alarm signal and / or the shutdown of the battery system can be initiated if the data from the battery monitoring system detects that individual battery cells 121 cannot be charged sufficiently and therefore a cell defect is suspected. List of reference symbols 10 passenger cabin 11 Wing 11a front wing 11b rear wing 11c upper wing 11d lower wing 11e Winglet 12 Cockpit 13 Main rotor 13a Rotor boom 14 rotor blades 15 propulsion propellers 15' first propulsion propeller 15" second drive propeller 15a Gondolas 16 Soundproofing ring 17 Vertical stabilizer 17a Rudder 18 Chassis 18a Nose wheel 18b tail wheel 20 generator 21 two-cylinder reciprocating piston engine 22 first cylinder-piston unit 23 second cylinder-piston unit 24 pistons 25 cylinders 26 Connecting rod 27 Crankshaft 27a Spur gear 28 timing belts 30 generators 30a balance weight 31 cam belt 32 camshaft 33 Valve 34 Oil pan 35 Oil pump 36 oil pump belt 37 oil filters 110 housings 111 Housing side panel 111a Overhang 111b Mounting hole 112 Housing front wall 120 cellblock 121 battery cells 122 Contact plate 123 Electronic board 124 Insulating sheath 125 connection module 125a Access opening 130 connection 131 electrical connection 132 pressure relief valve 133 Vacuum pump 140 cooling element S Focus

Claims

[1] Aircraft with at least one main rotor (13), at least one drive propeller (15) and a drive system, wherein the drive system a first electric motor which drives the main rotor (13) and draws electrical energy from a first drive battery, b. a second electric motor which drives the propulsion propeller (15) and draws electrical energy from a second propulsion battery, and c. a motor-generator unit comprising an internal combustion engine and at least one generator for generating electrical energy, which generator is electrically connected to the first drive battery and / or the second drive battery, wherein the internal combustion engine is a two-cylinder reciprocating piston engine with two cylinder-piston units in a tandem arrangement, wherein each cylinder-piston unit has a crankshaft and the crankshafts of both cylinder-piston units are mechanically coupled to one another, wherein at least one crankshaft, in particular both crankshafts each, is mechanically connected to a generator. [2] Aircraft according to claim 1, characterized by that the motor-generator unit has two generators, wherein a first generator is electrically connected to the first drive battery and a second generator is electrically connected to the second drive battery. [3] Aircraft according to claim 1 or 2, characterized bythat the drive system is a serial hybrid drive system. [4] Aircraft according to one of the preceding claims, characterized by that the combustion engine exclusively drives at least one generator or two generators. [5] Aircraft according to one of the preceding claims, characterized by that the internal combustion engine has a control system configured so that one cylinder-piston unit can continue to operate in the event of a failure of the other cylinder-piston unit. [6] Aircraft according to one of the preceding claims, characterized by that the drive system has at least one fuel tank, in particular for a methanol-based fuel. [7] Aircraft according to one of the preceding claims, characterized bythat the drive system has a coupling which is arranged between the main rotor (13) and the electric motor of the main rotor (13) in such a way that the main rotor (13) can be mechanically decoupled from the electric motor. [8] Aircraft according to one of the preceding claims, comprising a flight control system comprising electrically driven flight control rudders, the electrically driven main rotor (13) and two drive propellers (15), each driven by independent electric motors, wherein the electric motors of the flight control rudders, the main rotor (13) and the drive propellers (15) are connected to a control unit which has flight attitude sensors and is configured to control the individual electric motors such that the aircraft assumes a predetermined flight attitude, wherein the current flight attitude determined by means of the flight attitude sensors is compared with a predetermined desired flight attitude. [9] Aircraft according to claim 8, characterized by that the control unit has a data interface to a navigation system and is configured such that an autonomous flight from a predetermined starting location to a predetermined destination can be carried out using the data of the navigation system. [10] Aircraft according to one of the preceding claims, with the main rotor (13) having a plurality of adjustable rotor blades (14), wherein the main rotor (13) comprises at least one electromagnetic actuator coupled to the adjustable rotor blades (14) in order to set a predetermined adjustment angle of the rotor blades (14). [11] Aircraft according to claim 10, characterized by that the predetermined adjustment angle has three discrete adjustment angles. [12] Aircraft according to claim 11, characterized bythat a first adjustment angle is less than 0°, a second adjustment angle is at most 2.5° and a third adjustment angle is more than 2.5°. [13] Aircraft according to one of claims 10 to 12, characterized by that the adjustable rotor blades (14) are spring-loaded such that the adjustable rotor blades (14) assume a predetermined adjustment angle and are locked when the electromagnetic actuator is inactive. [14] Aircraft according to claim 13, characterized by that the adjustable rotor blades (14) assume and maintain the second adjustment angle. [15] Aircraft according to one of the preceding claims, with two drive propellers (15) which have parallel aligned axes of rotation, are each electrically driven and are encapsulated in a soundproofing ring (16). [16] Aircraft according to one of the preceding claims, characterized by - at least two counter-rotating propulsion propellers (15) for generating propulsion, which are arranged on opposite sides of a fuselage axis of the aircraft, wherein a first propulsion propeller (15') is driven by the second electric motor and a second propulsion propeller (15") is driven by a third electric motor, - at least two drive batteries, each of which is electrically connected to the first electric motor, the second electric motor and the third electric motor, wherein the first electric motor is adapted to temporarily drive the main rotor (13) and has an electrical power of at most 130 W / kg, in particular at most 125 W / kg, in particular at most 100 W / kg, in particular at most 90 W / kg, based on a take-off mass of the aircraft. [17] Aircraft according to claim 16, characterized bythat the first electric motor has an electrical power of at least 60 W / kg, in particular at least 70 W / kg, in particular at least 80 W / kg, relative to a take-off mass of the aircraft. [18] Aircraft according to claim 16 or 17, characterized by that a flight control is provided which is connected to a flight altitude sensor, a rotor blade adjustment of the main rotor (13) and the first, second and third electric motors in such a way that a. when the altitude sensor detects that the landing altitude falls below a predetermined first landing altitude, the main rotor (13) is driven by the first electric motor to provide lift support, the second and third electric motors being controlled in such a way that a torque of the main rotor (13) resulting from the lift support in the main rotor (13) is compensated, and b. if the altitude sensor detects that the aircraft falls below a predetermined second landing altitude, the rotor blade pitch is adjusted so that the aircraft's descent speed is reduced, in particular to a value of no more than 0.3 m / s. [19] A method for controlling an aircraft according to any one of the preceding claims for landing, comprising the following steps: a. detecting a drop below a predetermined first landing altitude by means of an altitude sensor; b. driving the main rotor (13) by means of the first electric motor when the first landing altitude is determined, wherein the second and third electric motors are simultaneously controlled in such a way that an additional torque generated by the first electric motor in the main rotor (13) is compensated; c. detecting a drop below a predetermined second landing altitude by means of the altitude sensor; d. Adjusting the rotor blade pitch when the aircraft falls below the second landing altitude so that the rate of descent of the aircraft is reduced, in particular to a value of no more than 0.3 m / s.

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