Method for assisting a pilot in controlling an aircraft, assistance system for an aircraft pilot, and aircraft

Abstract

Method for assisting a pilot (26) in controlling an aircraft (20) with multiple lift-generating propulsion devices (22, 23), the method comprising the following procedural steps: calculating instantaneous control limits (3, 4) for each control of the aircraft about its roll axis (x), control of the aircraft about its pitch axis (y), control of the aircraft about its yaw axis (z) and control of the total thrust T of all propulsion devices (22, 23), and displaying the calculated control limits (3, 4) by means of a display device (29).

Description

Technical FieldThe present invention relates to a method for assisting a pilot in the control of an aircraft comprising a plurality of lift-generating propulsion devices. The method essentially comprises a method step in which a plurality of instantaneous control limits are calculated and a further method step in which the calculated control limits are displayed by a display device.The invention also relates to an assistance system having a computing device and a display device, wherein the assistance system is preferably designed to carry out the aforementioned method.An aircraft with such an assistance system is likewise the subject matter of the present invention.Prior ArtAircraft with fly-by-wire control, particularly those with electric propulsion and vertical launch and landing mechanisms (eVTOL), often do not provide direct haptic feedback that allows the pilot to judge whether a requested control input can be performed at full or partial power (or not). Even if a control input can be fully executed, that is, if the aircraft can assume a flight condition demanded by the pilot, the remaining control margin may not be readily recognizable to the pilot under the given circumstances. In other words: the pilot may not be aware whether or which control reserves for controlling the aircraft still remain.The use of haptic feedback, in particular haptic feedback via the joystick of the aircraft, is known from the prior art. In this case, actuators bring about a force feedback which is transmitted from the joystick to the pilot.However, in eVTOL aircraft, in particular aircraft with a multi-rotor design and / or with a plurality of propulsion devices which have a fixed and / or adjustable angle of inclination or angle of attack for the individual propulsion devices, there is no clear independence between the control axes. For example, control input with respect to the roll or pitch axis in the hover mode may significantly reduce control margin along the yaw axis more than on the roll and pitch axes. This is due to the fact that such an aircraft can have N drive devices in the form of actuators (rotors or propellers together with the drive motor; N ∈ N), which can each be controlled by regulating the relevant rotor revolutions per minute (RPM)). The sum of the thrust force and torque generated by these N actuators and which can act on the aircraft body are used for controlling and stabilizing the aircraft. In this case, none of the drive devices can generate unlimited thrust and torque. There are also lower limits to the actuation, since a complete deactivation of a single actuator during flight is undesirable. While these restrictions for the individual actuators represent the relevant control limits from the standpoint of control theory, these restrictions are of little use for the commanding aircraft driver, especially since the saturation of an individual actuator does not mean that a control limitation of the aircraft as a whole has already been reached. This is due to the fact that the required workload can be distributed to all available actuators.Finally, the prior art force feedback may be ambiguous and confusion for the pilot in an eVTOL aircraft. This represents an incongruity in the man-machine interaction that cannot be tolerated in the light of flight safety.Independently of the aforementioned disadvantage, the introduction of active feedback by means of actuators inevitably increases the weight and complexity of the respective aircraft. In addition, the plurality of additional components increases the potential for possible sources of error.SUMMARY OF THE INVENTIONIt is an object of the present invention to satisfy the above-mentioned requirements from the prior art and / or to eliminate the disadvantages associated therewith. It is in particular an object of the present invention to specify a method, an assistance system and an aircraft in which the flight safety is increased within the scope of the interaction between the pilot and the aircraft without the total weight of the aircraft having to be increased.This object is achieved by a method according to claim 1. Advantageous further developments of the method according to the invention are the subject matter of the dependent claims 2 to 8 and / or are explained in the following description.To achieve the above object, the invention proposes a method suitable for assisting a pilot in the control of an aircraft having a plurality of lift-generating propulsion devices. In other words, the method is preferably applicable to aircraft having a plurality of propulsion devices, in particular a plurality of rotors. The drive devices are designed to generate lift in such a way that the aircraft can lift off and / or land, for example start perpendicularly and / or land perpendicularly. According to the angle of inclination of the propulsion devices with respect to the aircraft and / or the angle of inclination of the aircraft in space, the propulsion devices can additionally be designed to generate propulsion. The aircraft can be, for example, a multicopter, in particular an eVTOL, with a plurality of rotors. Preferably, each of the propulsion devices has an invariable angle of inclination with respect to the aircraft.According to the invention, the method comprises a method step in which instantaneous control limits of the aircraft are calculated. The calculation can be carried out, for example, by a calculation unit.The term "control limitation" is understood in the sense of the present invention to mean the physical limitation for the control of the aircraft. This means, for example, that the control of the aircraft outside the control limit is not physically possible and / or leads to the aircraft crashing. The control limits can correspond in particular to the flight states which can be physically reached by the aircraft. The control limits can be, for example, the limitation of a maximum roll angle with respect to the roll axis of the aircraft, a maximum pitch angle with respect to the pitch axis of the aircraft, a maximum yaw angle with respect to the yaw axis of the aircraft, and / or a maximum total thrust. Alternatively or additionally, the control limits can be designed as limits of the roll speed, the pitch speed and / or the yaw speed about the corresponding aircraft axis. Other limiting forms, for example the limiting of the maximum rotational acceleration and / or the maximum torques about the corresponding aircraft axis, are also possible as control limiting.The term "instantaneous" is understood in the sense of this invention to mean the state currently present in time. This means, for example, that the result of the current calculation can deviate from a result of the same calculation in the past and / or in the future, in particular in a past and / or future time step.According to the invention, the instantaneous control limits are calculated in each case with respect to a control of the aircraft about its roll axis, with respect to a control of the aircraft about its pitch axis, with respect to a control of the aircraft about its yaw axis, and with respect to a control of the total thrust of all drive devices. In this case, the physical situation is taken into account in that, in the case of an aircraft, in particular an eVTOL, having a plurality of lift-generating drive units, preferably in the case of a multicopter, all of the aforementioned control parameters influence one another or are dependent on one another. The control of the aircraft about its roll axis, about its pitch axis and / or its yaw axis can be dependent in particular on the arrangement and / or the instantaneous thrust power of the individual drive devices.In a further method step of the method according to the invention, the calculated control limits are displayed by a display device. The display can be effected, for example, in the form of a scale. By displaying the calculated control limits, these can be made accessible to the pilot by technical means.The method according to the invention advantageously has the effect that the pilot is constantly informed about the current limits or control limits of his aircraft by the aforementioned technical measures. This enables the pilot to control aircraft at all times within its physical capabilities. In other words, the method according to the invention prevents the aircraft from operating outside its physical options, which simultaneously reduces the risk of a potential crash. Accordingly, the method according to the invention has the advantage, among other things, that the flight safety is increased without the total weight of the aircraft having to be increased by additional components.In an exemplary embodiment of the method according to the invention, the control limits are calculated as a function of the current flight state of the aircraft. For example, the instantaneous position of the aircraft in the room and / or the incident flow of the aircraft can be detected by sensors for this purpose. Alternatively or additionally, the current control limits can be calculated as a function of the control of the aircraft.The term "control of the aircraft" can be, in particular, the control of the aircraft about its roll axis, about its pitch axis, about its yaw axis and / or the control of the total thrust of all drive devices. Independently of the foregoing, the instantaneous control limits may be calculated as a function of the pilot's instantaneous control inputs and / or as a function of the instantaneous control signals. The control signals may be, for example, signals from a flight control computer to the propulsion devices. In particular, the setpoint rotational speeds of the rotors and / or the actuation of the drive devices corresponding thereto can be control signals within the meaning of this invention.The aforementioned embodiment advantageously has the effect that all instantaneous control parameters are taken into account in the calculation of the instantaneous control limits. This has the advantage, among other things, that the physically present flight state of the aircraft can be detected particularly accurately with technical means, in order to thus increase the precision of the calculated control limits.In another exemplary embodiment, the control limits are calculated dynamically. In other words, the calculation of the control limits is repeated in terms of time. The control limits are preferably calculated periodically, i.e. repeatedly at predefined time intervals, for example. The time interval between the individual calculation steps is preferably selected to be so small that the calculation of the control limits takes place in real time.A dynamic, in particular periodic, calculation of the control limits advantageously has the effect that the calculated control limits always correspond to the current flight state of the aircraft. This has the advantage that the control limits provided by the method correspond sufficiently accurately to the currently physically maximum achievable flight state of the aircraft. In this way, the flight safety can be further increased.Preferably, each of the calculated control limits is displayed separately. In other words, each of the calculated control limits is displayed separately, in particular spaced apart, from the respective other calculated control limits. Such a display advantageously has the effect that the individual control limits can be detected particularly easily by the pilot. This can have the advantage that the flight safety is further increased. Alternatively, it is also conceivable for the calculated control limits to be displayed at least partially superimposed. This has the advantage that the display of the calculated control limits takes up or requires less display area. In this respect, in addition to the calculated control limits, further information can be displayed.In another exemplary embodiment, neutral control values may be displayed. The term "neutral control value" is understood in the sense of this invention to mean the neutral output value of the corresponding control parameter. For example, the neutral control value for controlling the aircraft about its roll axis is a roll angle of 0°. The same may also apply to neutral control values for the control of the aircraft about its pitch axis and / or yaw axis. Alternatively or additionally, the neutral control value for the control of the aircraft with respect to the corresponding aircraft axis can, however, also be a rotational speed of 0° / s, a rotational acceleration of 0° / s 2, and / or a torque of 0 Nm. The neutral control value with respect to the total thrust of all the driving devices may be 0 Newton (N). In other words, the drive devices generate no or a negligibly low lift in the presence of the neutral control value.Preferably, the calculated control limits are each displayed with respect to the respective neutral control value. For example, a control limit at which the control of the aircraft is not restricted at all or is only slightly restricted may be further spaced apart from the displayed neutral control value when displayed than a control limit at which the control of the aircraft is greatly restricted. Such a display can advantageously have the effect that the individual control limits can be detected particularly easily by the pilot. This can have the advantage that the flight safety is further increased.In a further exemplary embodiment of the method, for at least one control boundary, an area within the corresponding control boundary may be displayed visually differently than an area outside the corresponding control boundary. Preferably, the areas within each of the control limits are displayed visually differently with respect to the areas outside the corresponding control limits. The optical distinction can be realized, for example, by using different colors and / or different graphic structures. For example, the area within the control boundaries may be displayed in green color. The area outside the control limits can be displayed in black color. The optically differentiated representation can advantageously have the effect that the control ranges within the individual control limits can be detected by the pilot in a particularly simple manner. Accordingly, flight safety can be further increased thereby.At least one of the control limits may include an upper control limit and a lower control limit. Preferably, all control limits each comprise an upper control limit and a lower control limit.The upper control limit can be, in particular, a maximum control value. For example, the upper control limit for controlling the aircraft about its roll axis is a maximum roll angle that the aircraft can assume. The lower control limit preferably forms the maximum roll angle in the opposite direction. The same may also apply to the upper and lower control limits with respect to the pitch axis and / or yaw axis of the aircraft. Alternatively or additionally, the upper control limit with respect to the corresponding aircraft axis can, however, also be a maximum rotational speed in a first rotational direction, a maximum rotational acceleration in the first rotational direction, and / or a maximum torque in a first direction. The lower control limit may be a maximum rotational speed in a second rotational direction opposite the first rotational direction with respect to the corresponding aircraft axis. It is also conceivable that the lower control limit with respect to the corresponding aircraft axis is a maximum rotational acceleration in the second rotational direction and / or a maximum torque in a second direction. With respect to the total thrust of all the driving devices, the upper control limit may map the maximum thrust. In other words, the drive devices generate the maximum possible lift at the upper control limit.Preferably, for at least one control boundary, the upper and / or the lower control boundary is visually emphasized during the display. Particularly preferably, the respective upper and / or lower control limits are visually emphasized for all control limits. The control limits can be visually emphasized, for example, by signal colors, such as red, and / or a particularly conspicuous graphical representation, for example flashing. Additionally or alternatively, the upper and the lower control limits can also be displayed by displaying numerical values, in particular dynamically varying numerical values. The optical highlighting can advantageously have the effect that the corresponding control limits can be detected by the pilot in a particularly simple manner. Accordingly, flight safety can be further increased thereby.Independently of the above, the upper control limit and the lower control limit may be displayed as parallel lines. Alternatively or additionally, the upper control limit and the lower control limit may be indicated spaced with respect to the respective neutral control values. Preferably, the distance between the control limits and the neutral control value varies depending on the calculated current control limit.In a further exemplary embodiment of the method, an instantaneous control value is displayed in each case for controlling the aircraft about its roll axis, for controlling the aircraft about its pitch axis, for controlling the aircraft about its yaw axis and / or for controlling the total thrust of all drive devices.This may advantageously cause the pilot to be allowed to estimate its current control inputs with respect to the calculated control limits. In other words, such an embodiment notifies the pilot with technical means in which relation his control inputs stand to the physically possible flight control of the aircraft. This can have the advantage that flight safety is thereby further increased.In an advantageous development of the aforementioned embodiment, the representation of the instantaneous control value changes as soon as the corresponding instantaneous control value reaches the respectively calculated control limit. As a result, it can be optically signaled to the pilot by technical means that the pilot has exhausted the physically possible control potential of the aircraft with its control inputs.Another advantageous development of the method according to the invention provides that a graphic marker is used for displaying the instantaneous control value of the relevant component. The color of the marker, e.g. an arrow or triangle, is preferably changed depending on a distance of the associated value from the current boundaries. In this case, it can be provided that a display color of the marker corresponds to a display color of the corresponding control boundary when the current control boundary is reached, i.e. is changed to red, for example, while the display color is otherwise green, for example. When approaching the control limit, intermediate tones can also additionally be used, for example yellow and / or orange. Most preferably, when the instantaneous control limit is reached, an acoustic, visual and / or haptic alarm is additionally triggered. The marker itself can also be displayed flashing in this case.In an exemplary development of the method, the displays for the individual control limits can be arranged around a central display region. Said central display area can advantageously be used in a conventional manner for displaying instruments etc. of the aircraft. In particular, the so-called "basic T" known to the pilot, preferably comprising a display for artificial horizon, altimeter, travel meter and heading gyro, respectively, can be displayed there in order to make (also) the pilot available the cockpit design customary by other aircraft. As a result, the guidance of the aircraft can be further facilitated.In another exemplary embodiment of the method, the control of the aircraft may be described by a vector. The vector can correspond in particular to the current flight state of the aircraft. Preferably, the vector comprises four components, namely the three components of a torque, about the respective aircraft axes and the total thrust of all the drive devices. The three components comprising the torques may be dependent on the arrangement and / or the instantaneous thrust power of the individual drive devices. In order to calculate the control limits, the instantaneous limits of each of the four components of the vector are determined in particular.To calculate the vector by way of example, a three-dimensional torque vector τ (roll, pitch, and yaw moment in the unit Nm) acting on the aircraft body and the total thrust T (in the unit Newton) generated by all the drive devices in the rotor plane as a whole can be mathematically related to the individual actuator speed commands via the so-called motor matrix M:The vector RPM cmd; with RPM min ≤ RPM cmd ≤ RPM max may include the speed commands and preferably has the dimension N (N=number of propulsion devices), the engine matrix M is a 4×N matrix, and the vector γ has the dimension 4 (including the 3D torque τ and the total thrust force T). The square of the vector RPM cmd here means elementary squaring. The engine matrix M preferably depends on the position of the individual engines (or the respective rotors) in the aircraft and on the torque and thrust that can be generated by the engines / rotors. Note that these values may vary depending on various factors such as air density or (electric) voltage. As a rule, however, the motor matrix M can be approximated by a constant matrix for given conditions. In the following, it is assumed that a suitable matrix M is given or known.As a result of the lower and upper RPM limits RPM min, RPM max respectively, the set of all possible permutations of the vector γ is enclosed by a 4-dimensional convex hull, a so-called 4D polytop, which defines the control boundary of the aircraft. An advantageous embodiment of the invention now provides that this 4D polytope is approximated by a 4D hyperellipsoid with the best possible adaptation (e.g. according to the least squares method). For the purpose of illustration, reference is made to the description of the figures appended further below, in particular of FIG. 1.The restriction of vector γ to this 4D hyperellipsoid can be expressed as a quadratic inequality: with and the scalar values L max, M max, N max and T max, which indicate the half-axis length of the hyperellipsoid, and the scalar values L 0, M 0, N 0 and T 0, which define the center of the hyperellipsoid.It is important to note that the values for the maximum and center of the ellipsoidal axes are constant, while the current limits along each axis are highly dynamic and constantly change depending on the value of the currently (pilot) commanded 4D vector γ Cmd, i.e., depending on a current control input. A good example is a situation with a maximum thrust command that no longer leaves any clearance for roll and pitch commands or corresponding maneuvers, since all motors are already running at full speed and no rotational speed difference is available any longer in order to transmit a roll or pitch torque to the aircraft body.A possible aspect of the present invention is now to calculate the mentioned current control limits, e.g. on the basis of the hyperellipsoid representation explained above, and to use them for the display. In principle, the aim is thereby pursued to indicate dynamically to the pilot the currently still remaining control margin for each of the four available degrees of freedom (here corresponding to roll, pitch, yaw and lifting / total thrust). For a given γ Cmd to any epoch, i.e., in a given time interval, the control constraints for all axes may be calculated epoch-by-epoch in a time dependent manner. This dynamically calculated control limit can then be displayed to the pilot.The two individual limits (upper and lower control limits) for each of the axes are obtained by determining the two intersections of the line along the control axis by γ Cmd with the hyperellipsoid in the 4-dimensional torque-thrust space.Assuming that τ = γ - γ 0 the intersections to a general ratio κ of the ellipsoid along a search direction v are to be found. The following applies:This leads to solutions of the kindIf, in order to find all solutions, for the search directions v successively all unit directions (i.e. v=(1,0,0) T, (0,1,0,0)T etc.) are selected and additionally κ=1 is set, the calculations (with slight misuse of the notation) are simplified as follows:The dynamic upper and lower control limits, respectively, and for the four axes, i.e., the instantaneous control limit of vector γ with respect to each of the four components (axes of the ellipsoid) may be calculated as follows: where γ max= ( L max, M max, N max, T max)T and (.) 2 denote elementwise squaring, and where the square root is also elementwise performed. Subsequently, the dynamically determined control limit can be graphically displayed to the pilot as detailed above.The object of the invention, expressed at the beginning, is also achieved with an assistance system according to claim 9. Advantageous refinements of the assistance system according to the invention are explained in the following description.The assistance system according to the invention implements the same inventive idea as one of the aforementioned methods. The assistance system according to the invention is designed, in particular, to carry out one of the aforementioned methods. In this respect, all of the aforementioned features, combinations of features and advantages can also be combined with the assistance system or apply to the assistance system.The assistance system is designed in particular to assist a pilot in the control of an aircraft. The assistance system includes at least one computing device and a display device. The calculation device and the display device are preferably connected to one another in a signal-communicating manner. From the aforementioned independent, the calculation device and the display device can be spatially spaced apart from one another or can be formed within a common device.According to the invention, the calculation device is designed to calculate the instantaneous control limits for controlling the aircraft about its roll axis, controlling the aircraft about its pitch axis, controlling the aircraft about its yaw axis and controlling the total thrust of all drive devices. The computing device can be, for example, a processor and / or a computer.The display device is configured to display the calculated control limits, in particular the control limits calculated by the calculation device. The display device can be, for example, a display, in particular a head-up display (HUD) and / or a touchpad with display functionality.In an exemplary development of the assistance system according to the invention, the assistance system comprises an input device. The input device is preferably designed as a joystick (pilot control stick (PCS)). The input device can be connected and / or connectable to the computing device indirectly, i.e. directly, or directly, for example via one or more other devices, in a signal-communicating manner. Independently of the above, the input device can be connected and / or connectable to the display device in a signal-communicating manner indirectly, i.e. directly, or directly, for example via one or more other devices. The input device is preferably connected and / or connectable to the calculation device in a signal-communicating manner. The calculation device can in turn be connected and / or connected to the display device in a signal-communicating manner. The input device can be designed, for example, to record the pilot's control commands. Preferably, the pilot's control commands recorded by the input device are then provided to the calculation device as a control input, in particular as signals corresponding to the control input.The computing device may transmit the control input to the display device in modified and / or unmodified form. The calculation device can modify, in particular process and / or process, the control input, in particular within the scope of a regulation method. For example, the control method can have a pseudocontrol, so that the control method represents a type of mapping rule from control input and state to pseudocontrol.The object of the invention, expressed at the beginning, is also achieved with an aircraft according to claim 10. Advantageous refinements of the aircraft according to the invention are explained in the following description.The aircraft according to the invention implements the same inventive concept as one of the aforementioned methods and / or one of the aforementioned assistance systems. The aircraft according to the invention preferably has one of the abovementioned assistance systems. In this respect, all of the aforementioned features, combinations of features and advantages can also be realized / combined with or by the aircraft.According to the invention, the aircraft has, in addition to one of the abovementioned assistance systems, a plurality of lift-generating drive devices. At least one of the plurality of lift generating propulsion devices includes a rotor and an electric motor for driving the rotor. Preferably, all of the plurality of lift generating drive devices each include a rotor and an electric motor for driving the corresponding rotor.In an exemplary embodiment of the aircraft, the propulsion devices are arranged and / or configured on the aircraft such that the propulsion devices generate the lift necessary to maintain the aircraft in a substantially stationary hover position. In other words, the drive devices are arranged and / or configured on the aircraft in such a way that the drive devices can generate a vertical lifting force.The invention, as well as advantageous refinements of the invention, can / can also be described by way of example by the following aspects. The features listed in the aspects can be combined with the aforementioned features.Aspect 1 A method of assisting a pilot in controlling an aircraft having a plurality of lift generating propulsion devices, the method comprising the following method steps,Calculation of Current Control Limits for Eachcontrolling the aircraft about its roll axis,controlling the aircraft about its pitch axis,controlling the aircraft about its yaw axis; andcontrolling the total thrust T of all the driving devices, and displaying the calculated control limits by a display means.Aspect 2 Method according to Aspect 1, wherein the control limits are calculated as a function of the instantaneous flight state of the aircraft and / or the instantaneous control of the aircraft, preferably as a function of the control of the aircraft about its roll axis, about its pitch axis, about its yaw axis and the total thrust T of all drive devices.Aspect 3 Method according to one of the aforementioned aspects, wherein the control limits are calculated dynamically, preferably periodically.Aspect 4 Method according to one of the preceding aspects, wherein the control limits correspond to the flight states which can be physically reached by the aircraft.Aspect 5 The method of any preceding aspect, wherein each of the calculated control limits is displayed separately from the respective other calculated control limits.Aspect 6 The method of any preceding aspect, wherein at least one of the control limits is displayed relative to a neutral control value.Aspect 7 Method according to one of the preceding aspects, wherein for at least one control boundary, preferably for all control boundaries, an area within the corresponding control boundary is displayed visually differently than an area outside the corresponding control boundary.Aspect 8 The method according to any one of the preceding aspects, wherein at least one of the control limits, preferably all of the control limits, comprises an upper control limit and a lower control limit.Aspect 9 Method according to aspect 8, wherein for at least one control boundary, preferably for all control boundaries, the upper control boundary and / or the lower control boundary is graphically emphasized, preferably emphasized by a signal color.Aspect 10 The method of any of aspects 8 or 9, wherein the upper control limit and the lower control limit are depicted as parallel lines within the display panel.Aspect 11 Method according to one of the preceding aspects, wherein an instantaneous control value is displayed for the control of the aircraft about its roll axis, for the control of the aircraft about its pitch axis, for the control of the aircraft about its yaw axis and / or for the control of the total thrust T of all drive devices.Aspect 12 The method of Aspect 11, wherein the graphical representation of the instantaneous control value changes as the instantaneous control value reaches the respective control limit.Aspect 13 The method according to any of the preceding aspects, wherein the control of the aircraft is described by a vector, which vector comprises four components, namely the three components of a torque, about the respective aircraft axes and the total thrust, wherein the three components of a torque are dependent on the arrangement and the instantaneous thrust power of the individual drive devices, and the instantaneous limits of the vector with respect to each of the four components are determined for calculating the control limits.Aspect 14 An assistance system for assisting a pilot in the control of an aircraft, the assistance system comprising a calculation device for calculating instantaneous control limits for each of the control limitscontrolling the aircraft about its roll axis,controlling the aircraft about its pitch axis,controlling the aircraft about its yaw axis; andcontrolling the total thrust T of all the driving devices; anda display device for displaying the calculated control limits.Aspect 15 The assistance system according to aspect 14, wherein the assistance system is configured to carry out the method according to any of aspects 2 to 13.Aspect 16 Assistance system according to either of aspects 14 and 15, wherein the calculation device and the display device are connected to one another in a signal-communicating manner.Aspect 17 Assistance system according to one of aspects 14 to 16, wherein the assistance system has an input device, preferably a joystick, wherein the input device is connected to the calculation device in a signal-communicating manner and is designed to receive control commands of the pilot in order to provide the control commands of the pilot to the calculation device as control input.Aspect 18 An aircraft comprising an assistance system according to any of aspects 14 to 17, wherein the aircraft comprises a plurality of lift-generating propulsion devices.Aspect 19 The aircraft of Aspect 18, wherein each of the plurality of propulsion devices each includes a rotor and an electric motor for propelling the corresponding rotor.Aspect 20 The aircraft of any of aspects 18 or 19, wherein the propulsion devices are arranged such that the propulsion devices generate a vertical lift force.Brief Description of the DrawingsThe different and exemplary features described above can be combined with one another according to the invention, provided this is technically expedient and suitable. In this case, features disclosed in connection with a method can also be combined with device features. Further features, advantages and embodiments of the invention that can be combined are evident from the following description of exemplary embodiments and on the basis of the figures. The following are shown: FIG. 1 shows a (hyper) aircraft acceleration signal in a pseudo 3D representation in the torque-thrust space; FIG. 2 shows an exemplary embodiment of a dynamic representation of control limits, as can be generated with the aid of an exemplary embodiment of a method for assisting a pilot and / or an exemplary embodiment of an assistance system; FIG. 3 shows a representation analogous to FIG. 2 for two different points in time (epochs); FIG. 4 schematically shows a 2D diagram to illustrate the exemplary calculation of the control limits; and FIG. 5 shows an exemplary embodiment of an aircraft.WAYS OF CARRYING OUT THE INVENTIONFIG. 1 shows, by way of example, the ellipsoid E described further above for an aircraft in a pseudo 3D representation in the torque-thrust space, in which a dimension, in the present case the yaw torque ("yaw torque"), has been fixedly predefined. Accordingly, the ellipsoid E represents the currently available control volume in the remaining three spatial directions thrust ("thrust"), pitch torque ("pitch torque") and roll torque ("roll torque").The ellipsoid E describes the instantaneous control limits of a flight state vector γ (not shown in FIG. 1 ) which-starting from a center point of the ellipsoid E-can only lie on its surface or in its interior. The interior of the ellipsoid E indicates the physically accessible control range for the aircraft, i.e. those flight states which the pilot can currently realize with the aircraft.Based on this mathematical-physical circumstance, the pilot can now be clearly and easily depicted the basic situation according to FIG. 1 using technical means, preferably using an assistance system 1 and particularly preferably using a display device 29 (cf. FIG. 5 ). The technical contribution made thereby improves the man-machine interaction between pilot and aircraft in that the pilot knows the physical performance limits of the aircraft to be controlled at all times. A crash-risk oversteering of the aircraft by the pilot is thereby avoided.FIG. 2 shows a corresponding representation, as can be generated, for example, with the aid of the method according to the invention and / or the assistance system 1, in particular the display device 29 (cf. FIG. 5 ).In the context of the present invention, with a corresponding configuration, the pilot is dynamically indicated with the current control limits of the control volume for each component of the flight state vector γ (not shown in FIG. 2 ). For this purpose, a permanently predefined rectangular display field 2 is present on a display 1 of the assistance system for each component, that is to say for each control parameter of the aircraft in relation to the roll axis x, the pitch axis y and the yaw axis z (cf. FIG. 5 ), which is explicitly designated only for the roll moment. The displays or display regions 2 for the individual components are arranged around a central display region 2 a, in which central display region 2 acan be further displays relevant to control (not illustrated). Within a display panel 2, the dynamically calculated control value limits are displayed in red for the relevant component (at 3 and 4), one of these limits indicating the possible lower limit for the relevant component and the other limit the associated upper limit. Reference numeral 5 indicates the average value (average of the lower limit and the upper limit). The area 6 between the boundaries 3 and 4 is colored green and represents the physically accessible control area for the aircraft (with respect to the component in question).Reference numeral 8 denotes an instantaneous control value which indicates the currently controlled value of the relevant control limits 3, 4, in particular with respect to the upper control limit 3 and the lower control limit 4. The instantaneous control value 8 is white, but changes its color (limited instantaneous control value 9) as it approaches one of the control limits 3, 4. This is shown symbolically in FIG. 2 for the control with respect to the pitch axis ("pitch") and for the control with respect to the yaw axis ("yaw"). The instantaneous control value 8 is still far from the control limits with respect to the pitch axis ("pitch"). The limited instantaneous control value 9 with respect to the yaw axis ("yaw") has already reached the left-hand control limitation and is shown correspondingly differently in color.In other words, the yaw moment has reached its maximum (or minimum) permissible value; more control input γ Cmd in this direction is currently no longer possible. Accordingly, the relevant instantaneous control value 9 is colored differently in color, preferably red, and is located on the relevant control boundary.The thrust ("thrust") is almost maximum or near its possible maximum value. Accordingly, the instantaneous control value of the thrust is placed and also shown visually differently. In particular, the framework of the instantaneous control value is somewhat thicker and the coloring can also be selected differently. For example, the instantaneous control value can be colored yellow.The instantaneous control values for rolls ("roll") and pitch ("pitch") are colored in the vicinity of the neutral control value 5 and therefore white.The control limits 3 and 4 and also the region 6 within the control limit 3, 4 together with the neutral control value 5 are dynamically displaceable within the (permanently predefined) display fields 2, depending on the current flight state or pilot control command. In this case, the display within the display panels 2 does not have to be symmetrical with respect to the panel boundaries (rectangular border), in particular for the push.The following special facts in FIG. 2 are additionally explicitly identified there:The instantaneous control value 8 denotes a current commanded pitch torque (element of the 4x1 vector y; not shown) on the aircraft fuselage or its display, which-like all other components also-is provided at a preferably constant update rate, in particular by the mentioned calculation unit (cf. FIG. 5 ), e.g. by a flight control law implemented there. The commanded roll, yaw and thrust value (also elements of vector y) is displayed in the other three display fields 2, as already described.At the limited instantaneous control values 9, none of the four commanded torque / thrust values (elements of vector y) can by definition exceed the hyperellipse of the control volume and thus the dynamic limits (cf. also reference numerals 3, 4), as is exemplarily shown for the yaw value in FIG. 2.The thick, preferably red, bars or lines indicate the current control limit for the thrust, as indicated by reference numeral 10. These beams (corresponding to the limits 3, 4) already mentioned move continuously depending on the state of the aircraft and the pilot's inputs and thus increase and decrease the control volume on the respective axis. This applies analogously also to the displays for roll, pitch and yaw.This can also be seen in FIG. 3, in which the thick bars for the upper control limit 3 and the lower control limit 4 change dynamically between an epoch (time point) k and a later epoch (time point) k+1. It can be seen, for example, that the control volume is significantly smaller in epoch k+1 (on the right in FIG. 3 ) in particular on the yaw axis than in epoch k (on the left) due to the large roll and pitch commands.The available control volume is preferably displayed in green (see hatched area described by reference numeral 11 in Fig. 2). The corresponding region related to the control about the roll axis of the aircraft (cf. reference numeral 6) changes its size dynamically depending on a position of the boundary beams (boundaries 3, 4).FIG. 4 shows a 2D representation to illustrate the calculation of the dynamic limits, as described mathematically in detail further above.Only the thrust (Y axis) and the pitching moment (X axis) are shown here by way of example. The scalar values N 0 and T 0 indicate the center of the hyperellipsoid located at the intersection of the two associated lines. The scalar values N max and T max describe the associated half-axis length of the hyperellipsoid E. The same applies to the other axes (roll and pitch) not shown here.Depending on the position of vector γ, dynamic upper and lower control limits result for the individual components of vector γ, referred to as and in FIG. 4. These also depend for each individual component on the values of the respective other components, as already described.For example, the instantaneous control limit and for the thrust result by running from γ parallel to the ordinate up to the points of intersection with the ellipsoid E. The same applies to the pitching moment and its instantaneous control limitation with respect to the abscissa (and correspondingly to the further components or dimensions; not shown in FIG. 4 ).FIG. 5 schematically shows an exemplary embodiment of an aircraft 20 in the form of an eVTOL. The aircraft 20 has a plurality of propulsion devices 21, each of the propulsion devices 21 has a fixed pitch propeller or rotor 22 and an electric motor 23 for driving the rotor 22, the rotors 22 being all arranged in or parallel to a common plane 24, the rotor blades being inclined with respect to the plane 24 or with respect to a supporting structure (fixed pitch angle).Below the plane 24 is the aircraft cell 25 together with the cockpit. A pilot 26 flies or controls the aircraft 20 with the aid of an input device 27 in the form of a joystick 27. the aircraft 20 further comprises a calculation device 28. the calculation device 28 is connected to the input device 27 in a signal-communicating manner. Reference numeral 29 denotes a display device which, in the exemplary embodiment illustrated in FIG. 5, is embodied as a display (screen) 29. The display device 29 is in turn connected in signal communication with the calculation device 28. The computing device 28, the input device 27 and the display device 29 together form an assistance system 1.If now the pilot 26 moves the joystick 27 in order to control the aircraft 20, the calculation device 28 receives the corresponding control signals and can carry out the calculations provided within the scope of the method described above. Thus, the dynamically calculated current control limits can subsequently be displayed to the pilot 26 via the display 29.List of reference characters1 Assistance system 2 Display area 3 Upper control limit 4 Lower control limit 5 Neutral control value 6 Area within the control limits 7- 8 Instantaneous control value 9 Limited instantaneous control value 20 Aircraft 21 Drive device 22 Rotor 23 Motor / electric motor 24 Plane 25 Aircraft cell 26 Pilot 27 Input device / joystick 28 Calculation device 29 Display device E Ellipsoid γ Vector T Thrust X Rolling axis of the aircraft Y Pitch axis of the aircraft Z Yaw axis of the aircraft

Claims

A method for assisting a pilot (26) in controlling an aircraft (20) having a plurality of lift generating propulsion devices (21), the method comprising the following method steps, (a) calculating instantaneous control limits (3, 4) for each of (a1) a control of the aircraft about its roll axis (x), (a2) a control of the aircraft about its pitch axis (y), (a3) a control of the aircraft about its yaw axis (z) and (a4) a control of the total thrust T of all propulsion devices (21), and (b) displaying the calculated control limits (3, 4) by a display device (29).Method according to Claim 1, wherein the control limits (3, 4) are calculated as a function of the instantaneous flight state of the aircraft (20) and / or the instantaneous control of the aircraft (20), preferably as a function of the control of the aircraft about its roll axis (x), about its pitch axis (y), about its yaw axis (z) and the total thrust T of all the drive devices (21).Method according to one of the preceding claims, wherein the control limits (3, 4) are calculated dynamically, preferably periodically.Method according to one of the preceding claims, wherein at least one of the control limits (3, 4) is displayed relative to a neutral control value (5).Method according to one of the preceding claims, wherein at least one of the control limits (3, 4), preferably all of the control limits (3, 4), has an upper control limit (3) and a lower control limit (4).Method according to claim 5, wherein for at least one control boundary (3, 4), preferably for all control boundaries (3, 4), the upper control boundary (3) and / or the lower control boundary (4) is displayed graphically highlighted, preferably highlighted by a signal colour.Method according to one of the preceding claims, wherein an instantaneous control value (8) is displayed for controlling the aircraft about its roll axis (x), for controlling the aircraft about its pitch axis (y), for controlling the aircraft about its yaw axis (z) and / or for controlling the total thrust (T) of all drive devices (22, 23).Method according to any one of the preceding claims, wherein the control of the aircraft is described by a vector (γ), which vector (γ) comprises four components, namely the three components of a torque (τ) about the respective aircraft axes (x, y, z) and the total thrust (T), wherein the three components of a torque (τ) are dependent on the arrangement and the instantaneous thrust power of the individual drive devices (22, 23), and for calculation of the control limits (3, 4) the instantaneous limits of the vector γ with respect to each of the four components is determined.Assistance system (1) for assisting a pilot (26) in the control of an aircraft (20), the assistance system (1) comprising (a) a calculation device (28) for calculating instantaneous control limits (3, 4) for respectively (a1) a control of the aircraft about its roll axis (x), (a2) a control of the aircraft about its pitch axis (y), (a3) a control of the aircraft about its yaw axis (z) and (a4) a control of the total thrust T of all drive devices (22, 23), and (b) a display device (29) for displaying the calculated control limits (3, 4).Aircraft (20) with an assistance system (1) according to claim 9, wherein the aircraft (20) comprises a plurality of lift-generating propulsion devices (21).

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