DE102020209359A1 - Short or vertical takeoff capable aircraft, power electronics and method of operating the aircraft - Google Patents

Abstract

The invention relates to an aircraft (10) capable of short or vertical takeoff, having a plurality of drive units (24), each having at least one rotor or propeller (28) and an electric motor (30), the electric motors (30) being powered by at least a first energy unit ( 36) coupled and the drive power of the electric motors (30) can be regulated by means of power electronics (34).

Description

technical field

The invention relates to an aircraft capable of short (STOL) or vertical takeoff (VTOL) as it can be used, for example, as a drone or for transporting people or loads with a payload of up to several 100 kg. Typical representatives of such aircraft are multicopters or lift-and-cruise or tilt-wing aircraft. Furthermore, the invention relates to power electronics as part of the aircraft according to the invention and to a method for operating the aircraft.

State of the art

An aircraft with the features of the preamble of claim 1 is, for example, in the form of a drone from the CN 204548495 U known. The known aircraft has several drive units, each arranged on support arms, each with an electric motor and a rotor or propeller. By controlling the drive power and the position of the drive units on the aircraft, the desired flight maneuvers, ie takeoff, landing and a cruising flight between takeoff and landing, can be carried out with a route profile.

Such an aircraft typically has at least one first energy unit in the form of a rechargeable battery or an accumulator, which is arranged centrally and is used to supply the individual drive units. In this case, the size or the power output of the at least one first energy unit is typically designed for the starting process (with corresponding power reserves), since that is where the highest energy output is required from the drive units for vertical takeoff.

Disclosure of Invention

The short or vertical takeoff capable aircraft with the features of claim 1 has the advantage that it enables particularly safe flight operations. In particular, the aircraft according to the invention makes it possible, for example, to land the aircraft safely in the event of damage or even failure of the at least one first energy unit. The aircraft according to the invention also makes it possible to design the at least one first energy unit to be particularly simple or light and therefore also inexpensive, since it does not necessarily have to be designed to be able to carry out all flight states of the aircraft solely by its energy output.

The above-mentioned advantages of the aircraft are achieved according to the invention in that at least one second energy unit is provided in the form of a rechargeable battery, which interacts directly or indirectly with the electric motors of the drive units, and that the at least one second energy unit is designed to temporarily additionally the at least one first energy unit to interact with the drive units.

Such a design of the aircraft with at least one second energy unit thus makes it possible, for example for take-off and landing phases in which a particularly high energy requirement is required for an aircraft taking off or landing vertically, to cover the required energy requirement by switching on the at least one second energy unit , especially in the event that the at least one first energy unit cannot perform these flight tasks in terms of their power output.

The redundant or at least two-fold energy unit (implemented by the at least one first and the at least one second energy unit) also enables a particularly safe transport of goods or people insofar as unsafe flight conditions or crashes of the aircraft are triggered if the at least one first energy unit fails can be avoided by appropriate design of the second energy unit, which is also advantageous for the transport of dangerous goods or increases the acceptance of (person) transport with such an aircraft. When designing both the power electronics for drive control and the electric motors of the drive units, care must be taken that these are designed for simultaneous loading by the at least two energy units.

Advantageous further developments of the aircraft capable of short or vertical takeoff are listed in the dependent claims.

A particularly preferred structural design of the aircraft provides that an adjusting device is provided for setting a rotor or propeller pitch. This allows flight conditions to be optimized, for example with regard to their noise emissions or the lowest possible energy consumption. For example, high rotor or propeller pitches can be set to achieve relatively high cruising speeds of the aircraft, while the highest possible thrust is required for the take-off and landing phase, as is typically achieved here by a relatively small rotor or propeller pitch.

In order to enable particularly reliable operation of the aircraft on the one hand and to minimize the energy losses during the transmission of energy from the at least one second energy unit to the electric motor on the other hand, it is provided that each drive unit is assigned a separate second energy unit and that the second energy unit is in the area or is arranged near the electric motor and/or the power electronics.

In addition, it is particularly preferred if the energy content or the power output of the at least one second energy unit is designed to carry out a landing maneuver of the aircraft solely using the at least one second energy unit. In other words, this means that failure of the at least one first energy unit does not lead to damage or to a crash of the aircraft.

In order in particular to avoid complex charging infrastructures for the at least one second energy unit, it is also advantageous if the at least one second energy unit can be charged by the at least one first energy unit. This typically takes place in phases in which the at least one first energy unit can provide power reserves in order to charge the at least one second energy unit. Alternatively, however, solutions are also conceivable in which, for example, the at least one second energy unit is inductively charged when the aircraft is at a take-off or landing site provided for this purpose. Inductive charging has the particular advantage that no cable connections or the like are required or have to be established between a charging device or an energy supply and the at least one second energy unit.

With regard to the at least one first energy unit, there are different options that are based in particular on the specific requirements of the aircraft, such as its required range, payload, size, etc. In particular, it is conceivable that the at least one first energy unit is designed as a rechargeable battery, as a hybrid drive using an internal combustion engine, or as a fuel cell system.

It is very particularly preferred in connection with the adjusting device for adjusting the pitch of the rotor or propeller if a flight computer device is provided which interacts with the adjusting device and a powertrain control unit and is designed to adapt the speed and the pitch of the rotors or propeller to predetermined criteria , in particular to adapt a (maximum) noise emission and/or a flight route profile.

The invention also includes power electronics as part of an aircraft according to the invention that has been described so far. The power electronics are characterized in that they are designed to control at least the speed or power of at least one electric motor of the drive units and are connected to the at least one first and at least one second energy unit.

Finally, the invention also includes a method for operating an aircraft according to the invention. The method is characterized in that the at least one second energy unit is activated to support the at least one first energy unit when a power requirement of the drive units cannot be covered by the at least one first energy unit alone.

In a development of the method, it is provided that the at least one second energy unit is activated during takeoff or landing of the aircraft. This enables either increased power reserves, for example for gust compensation or the like in particularly critical flight phases, or it is possible to build the at least one first energy unit particularly small or compact in terms of its size or power output and thus also in terms of its weight and investment costs .

Further advantages, features and details of the invention result from the following description of preferred embodiments of the invention and from the drawings.

character list

• 1 shows a highly simplified perspective representation of a vertically take-off aircraft in the form of a drone,
• 2 a diagram to clarify the essential components of the aircraft according to the 1 and
• 3 a flight profile during a deployment of the aircraft according to the 1 in a visual representation.

Embodiments of the invention

Identical elements or elements with the same function are provided with the same reference numbers in the figures.

In the 1 is greatly simplified a short or vertical takeoff capable aircraft 10 in form a drone 12 shown. The (unmanned) drone 12 can, for example, be designed for industrial purposes for taking aerial photographs or for transporting goods, for example as part of a package delivery. However, the invention should not be limited to such (relatively small) drones 12 . Rather, the aircraft 10 can also be, for example, a manned aircraft 10 that can fly autonomously. Such aircraft 10 can, for example, connect inner-city areas with an airport and be used as a feeder service for passenger transport. In this case, the aircraft 10 can typically have a mass of up to 1,000 kg or more.

The aircraft 10 or the drone 12 has a central load or device area 14, shown cylindrically in the exemplary embodiment for the sake of simplicity, from which three extension arms 18 protrude at equal angular distances about a vertical axis 16, each of which is about a longitudinal axis 20 by means of not-shown Adjustment devices are pivotable. On the side facing away from the device area 14 , identically designed drive units 24 are arranged on the cantilever arms 18 .

Each of the drive units 24 has, for example, a sleeve-shaped jacket 26 connected to the cantilever arm 18, within which one of a in the 1 not recognizable electric motor driven propeller 28 is arranged.

In addition, it is mentioned that it is of course within the scope of the invention to provide more than three drive units 24, for example four to six drive units 24. It is also conceivable for each drive unit 24 to be assigned two propellers 28 arranged coaxially to one another. It is also not necessary for the drive units 24 to have propellers 28 arranged inside a casing 26 . Rather, the propellers 28 or rotors can also be arranged to rotate freely.

The flight maneuvers of the drone 12 or the aircraft 10 are effected in a manner known per se by controlling the power of the propellers 28, i.e. in particular by influencing their speed n, and by pivoting the extension arms 18 on the device area 14, for example in order to transition to a vertical take-off position in a cruise flight where the aircraft 10 can cover a horizontal distance.

In the representation of 2 In addition to the propeller 28, a boostable electric motor 30 for driving the propeller 28 at least indirectly can be seen. The electric motor 30 acts on the propeller 28 via an adjustment device 32 . The adjustment device 32 serves to set a (propeller) pitch s of the propeller 28 . To achieve a high horizontal speed or a high jet speed of the propeller 28, it is typically necessary to set a relatively high pitch s on the propeller 28, while in (vertical) takeoff or landing phases, in contrast, smaller pitches s are typically required or desired.

Furthermore, it is noted that the noise emission of the propeller 28 also depends on its adjusted pitch s, such that a high difference between the jet velocity of the propeller 28 and the surrounding air typically leads to higher noise emissions.

Each of the electric motors 30 is assigned, for example, power electronics 34 which serve to control the power or speed n on the electric motor 30 . The electric motor 30 is supplied with energy via the power electronics 34 via at least one first energy unit 36, which can be designed in the form of a (rechargeable) battery, a hybrid drive using an internal combustion engine or as a fuel cell system.

Typically, the design of the aircraft 10 is such that the at least one first energy unit 36 simultaneously serves to supply energy to all drive units 24 and is arranged in the equipment area 14 .

The control of the power electronics 34 and the adjustment device 32, i.e. the setting of the speed n of the electric motor 30 and the pitch s of the propeller 28 is also carried out by means of an engine train control unit 40, which in turn interacts with a flight computer device 42 or can be controlled by it.

In addition to the at least one first energy unit 36 , the aircraft 10 has at least one second energy unit 44 in the form of a rechargeable battery 45 . The at least one second energy unit 44 for supplying energy to the drive unit 24 or the electric motor 30 can be controlled by the powertrain control unit 40 in order to enable a maximum torque mode in the boost mode of the electric motor 30 or, in the event of a failure/defect in the first energy unit 36, operation via the second energy unit 44 to realize. There is also a connection between the at least one second energy unit 44 and the electric motor 30 via the power electronics 34.

A charging connection 46 is preferably provided between the two energy units 36 and 44, which enables the second energy unit 44 to be charged by the first energy unit 36, in particular in operating states of the aircraft 10 in which the at least one first energy unit 36 can supply more energy than is required for the safe operation of the aircraft 10. In addition, it can also be provided that the at least one first energy supply 36 is designed in such a way that it is not sufficient on its own to enable all flight states of the aircraft 10 that are usually provided. In particular, takeoff and landing phases of the aircraft 10 are typically associated with an increased energy requirement. In this case, the drive units 24 can be supplied with energy in addition to the at least one first energy unit 36 via the at least one second energy unit 44, which is also preferably arranged in or near the area of the drive units 24 to reduce losses.

In the 3 Graph F of a typical flight profile of the drone 12 during an operation, for example when delivering a parcel, is shown. The (horizontal) flight path s of the drone 12 is shown in the horizontal direction, while the flight altitude h of the drone 12 is shown in the vertical direction. The flight profile is characterized by a starting phase SP, during which the drone 12 reaches a cruising altitude hi. The starting phase SP is followed by a cruising phase RP, which includes two sections a and b. The travel phase RP is in turn followed by a descent phase AP and a landing phase LP.

During the take-off phase SP and the landing phase LP, in which an increased power requirement of the drone 12 is typically required to drive the drive units 24, it can be provided that in addition to the at least one first energy unit 36 the at least one second energy unit 44 supplies the energy or provides electrical power to the drive units 24 . In contrast, it can be provided that in the cruising flight phase RP the power supply to the drive units 24 is provided exclusively by the at least one first power unit 36 .

In addition, it is mentioned that during the individual flight phases, the flight computer device 42 on the drive units 24 or the electric motors 30 sets the respectively optimum speed n or pitch s on the propeller 28 . For example, it can be provided that when flying over a nature reserve in section a during the cruising flight phase RP, low noise emissions are desired, which, however, require increased energy. For this purpose, a relatively small gradient s on the propeller 28 can be advantageous, for example, at a relatively high speed n of the electric motor 30 . In contrast, section b of the cruise flight phase RP, in which no noise emissions are relevant, can be about flight operations that are as economical as possible, i.e. optimized in terms of efficiency or minimized energy consumption. In this case, provision can be made, for example, for the speed n of the electric motors 30 to be reduced, while the pitch s on the propellers 28 is increased. For example, a constant flight speed can also be achieved during the two sections a and b during the cruising flight phase RP.

• So kann es beispielsweise vorgesehen sein, dass für verschiedene Kriterien wie eine maximale Energieeffizienz, eine minimale Schallemission oder ein maximaler Schub Idealwerte für die Drehzahl n und die Steigung s abhängig vom jeweiligen Sollschub und der durchströmenden Luftgeschwindigkeit oder der Fluggeschwindigkeit gesteuert vorgegeben werden. Diese Steuerwerte für die Drehzahl n und die Steigung s können auch für eine unterschiedliche Gewichtung der angesprochenen Kriterien vorgegeben werden, beispielsweise ein Faktor von 0,7 für die Schallemission + ein Faktor von 0,3 für die Energieeffizienz. Auch kann es vorgesehen sein, dass die Steuerwerte beim Auftreten von beispielsweise Störungen oder Verschleiß angepasst bzw. verändert werden, um die Auswirkungen zu minimieren. Als mögliche Regler zur Optimierung der Kriterien können vorzugsweise Extremwertregler, hilfsweise lineare oder nichtlineare Optimierer eingesetzt werden. Die Schrittweite der Optimierer bzw. Regler kann bei Annäherung an das jeweilige Optimum verringert werden.

Provision can also be made for the flight computer device 42, in conjunction with the engine train control unit 40 and the power electronics 34, to be designed to optimize the flight operation of the drone 12 in various ways on the basis of the following explanations, individually or in combination with the measures addressed :

• For example, it can be provided that for various criteria such as maximum energy efficiency, minimum noise emission or maximum thrust, ideal values for the speed n and the gradient s are specified in a controlled manner depending on the respective target thrust and the air speed flowing through or the flight speed. These control values for the speed n and the gradient s can also be specified for a different weighting of the criteria mentioned, for example a factor of 0.7 for noise emission+a factor of 0.3 for energy efficiency. Provision can also be made for the control values to be adapted or changed when, for example, faults or wear occur, in order to minimize the effects. Extreme value controllers, alternatively linear or non-linear optimizers, can preferably be used as possible controllers for optimizing the criteria. The increment of the optimizers or controllers can be reduced when approaching the respective optimum.

The regulation can be designed in such a way that the speed n and the gradient s are set in such a way that their quotient for the maximum energy efficiency is at a maximum. The regulation can also provide for an adjustment of speed n and slope s, at which a specified thrust is achieved with minimal noise emissions. This minimum sound emission can also be filtered or evaluated with a view to a sensitivity-frequency curve of human hearing, so that the minimum sound emission is one for the human hearing is minimal sound emission. With a view to the noise emission or its minimization, it can also be provided that the speeds n of the individual propellers 28 are regulated in such a way that noise beats, which are perceived as unpleasant, are avoided.

In order to achieve maximum thrust, for example for takeoff or landing, the controller can set an optimum for this from the corresponding speed n, gradient s and motor power of the electric motors 30 . For this purpose, provision can also be made for electric motors 30 to be able to output increased power, for example by switching from sinusoidal commutation to block commutation, or by connecting the at least two energy units 36 and 44 in series.

The aircraft 10 described so far can be altered or modified in many ways without departing from the spirit of the invention. For example, it can be provided that the at least two energy units 36, 44 and/or the electric motors 30 of the drive units 24 are equipped with cooling devices in order to avoid thermal overloading of the components mentioned.

QUOTES INCLUDED IN DESCRIPTION

This list of documents cited by the applicant was generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Patent Literature Cited

• CN 204548495 U [0002]

Claims (12)

Aircraft (10) capable of short or vertical takeoff, with a plurality of drive units (24), each having at least one rotor or propeller (28) and an electric motor (30), the electric motors (30) being coupled to at least one first energy unit (36) and the drive power of the electric motors (30) can be regulated by means of power electronics (34), characterized in that at least one second energy unit (44) is provided in the form of a rechargeable battery (45) which is connected directly or indirectly to the electric motors (30) of the Drive units (24) interacts, and that the at least one second energy unit (44) is designed to temporarily interact with the drive units (24) in addition to the at least one first energy unit (36). aircraft after claim 1 , characterized in that an adjustment device (32) for setting a rotor or propeller pitch (s) is provided. aircraft after claim 1 or 2 , characterized in that each drive unit (24) is assigned a separate second energy unit (44), and that the second energy unit (44) is arranged in the area or near the electric motor (30) and/or the power electronics (34). Aircraft after one of Claims 1 until 3 , characterized in that the energy content or the power output of the at least one second energy unit (44) is designed to perform a landing maneuver of the aircraft (10) without the at least one first energy unit (36). Aircraft after one of Claims 1 until 4 , characterized in that the at least one second energy unit (44) can be charged by the at least one first energy unit (36). Aircraft after one of Claims 1 until 5 , characterized in that the at least one first energy unit (36) is designed as a rechargeable battery, as a hybrid drive using an internal combustion engine or as a fuel cell system. Aircraft after one of claims 2 until 6 , characterized in that a flight computer device (42) which interacts at least indirectly with the adjustment device (32) and a power train control unit (40) is provided and is designed to calculate the speed (n) and the pitch (s) of the rotors or propellers ( 28) to adapt to specified criteria, in particular noise emissions and/or a flight route profile. Power electronics (34), formed as part of a according to one of Claims 1 until 7 trained aircraft (10), characterized in that the power electronics (34) for controlling at least the speed (n) or power of at least one electric motor (30) of the drive units (24) and formed with the at least one first and at least one second energy unit ( 36, 44). Method for operating an aircraft (10) according to one of Claims 1 until 7 is designed, characterized in that the at least one second energy unit (44) is activated to support the at least one first energy unit (36) if a power requirement of the drive units (24) cannot be covered exclusively by the at least one first energy unit (36). . procedure after claim 9 , characterized in that the at least one second energy unit (44) is activated during a take-off phase (SP) and/or a landing phase (LP) of the aircraft (10). procedure after claim 9 or 10 , characterized in that the speed (n) and the pitch (s) of the rotors or propellers (28) are set based on predetermined criteria, such as maximum thrust, minimum noise emissions or maximum energy efficiency. procedure after claim 11 , characterized in that for the simultaneous fulfillment of several criteria these are weighted differently and the speed (n) and the pitch (s) of the rotors or propellers (28) are each optimized to achieve the weighted criteria.

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