rexresearch1
Michael WINDERL, et al.
CycloRotor
CycloRotor
https://www.cyclotech.at/rotor/
CycloRotor Aviation Propulsion
CycloRotors provide unmatched manueverability by directing thrust 360° around their rotational axis, surpassing traditional one-direction systems. CycloTech’s 5th-generation CycloRotors powered an entire aircraft in 2021, demonstrating revolutionary design flexibility for UAVs and UAM vehicles.
360° Thrust Vectoring System
A CycloRotor is a propulsion unit that can change the magnitude and direction of thrust without the need to tilt any aircraft structures. It contains several parallel blades rotating around a central rotation axis. The thrust is generated by a combined airflow through the rotor originating from each blade and its periodic change of the pitch angle during one rotation. The individual pitch angle of the blades is controlled by a certain pitch mechanism. Each blade is mechanically connected to a central hub with a conrod. The Cyclogyro-Rotor magnitude of thrust and its direction can be directly controlled by the eccentric positioning of this hub. This enables an easy and fast way of thrust-vector control of the propulsion unit. CycloRotors enable an immediate thrust generation 360° around the rotation axis – at constant rotation speed and direction – within fractions of a second.
Unique Characteristics of CycloRotors
360° THRUST VECTORING ... Easy transition hover to forward flight ... Superior manoeuvrability and gust control ... Decoupling of flight path and vehicle attitude ... COMPACT DESIGN .. Small footprint ... CONFINED AREAS ... CROWDED AIRSPACE ... ADVERSE WEATHER ... ADDED SAFETY ... AGILITY ... COMFORT ... CycloRotor Development
The CycloRotor’s innovative propulsion system requires an ultra-light yet durable design, utilising advanced materials and manufacturing techniques. CycloTech has created an automated toolchain combining CFD simulations, multi-body dynamics, and FE modelling to optimise performance, with rigorous testing in wind tunnels and flight. In just three years, efficiency has more than doubled, paving the way for even greater advancements in this evolving technology.
CycloRotor CR-42*
The current rotor design with a one-sided mounting offers an easy connection to any aircraft structure via a simple suspension system. CR42 is a fully electric CycloRotor with an electric motor and a gearbox as a drivetrain. The core of the agility is the patented mechanical configuration and control system.
https://www.youtube.com/watch?v=DGUgpyniLsk
Discover CycloTech #4 | What is the CycloRotor and why it´s a game changer for aviation
https://www.youtube.com/watch?v=oYOZYsHs7_Q
https://www.youtube.com/watch?v=qKgskY_69XE
https://www.youtube.com/watch?v=YYCNMXIyyk8
US11479356 -- Driver device for an aircraft [ PDF ]
Inventor(s): WINDERL MICHAEL [DE]; LANSER STEPHAN [AT] +
Applicant(s): CYCLOTECH GMBH [AT] +
The invention relates to a propulsion device for an aircraft, comprising a blade (2) which can be rotated about an axis of rotation (51) of the propulsion device along a circular path (52) and is mounted for pivoting about a blade bearing axis parallel to the axis of rotation; a pitch mechanism having a coupling device (31) and a bearing device (33); and an offset device (4) to which the blade is coupled, the offset device defining an eccentric bearing axis (41) which is mounted at an adjustable offset distance. The coupling device is coupled to the blade at a coupling point (32) which is positioned in such a way that the plane that comprises the blade bearing axis and the coupling point and the tangential plane to the circular path through the blade bearing axis include a certain, non-vanishing angle (wα) when the offset distance is set to zero. According to a second aspect the blade bearing axis is shifted toward the axis of rotation by a certain distance relative to the plane that extends through the center of mass of the blade and that extends parallel to the axis of rotation and to the chord of the blade.
This application is a national stage entry of PCT/EP18/84371, filed Dec. 11, 2018 and entitled Driver Device for an Aircraft, which claims priority to German application number 102017011890.6, filed Dec. 14, 2017 and entitled Driver Device for an Aircraft. Each of these applications is incorporated by reference in its entirety.
The subject matter of the invention is a propulsion device for an aircraft. Specifically, the invention relates to a cyclogyro rotor with reduced loads on the structural elements of the rotor.
A cyclogyro rotor is based on the principle of thrust generation with rotating blades. In contrast to classical rotating blades, such as those which are used in the propulsion device of a helicopter, the axis of rotation of the blades of a cyclogyro rotor is oriented parallel to the longitudinal axis of the blades. The direction of thrust of the entire cyclogyro rotor is normal to the axis of rotation.
Cyclogyros are aircrafts using cyclogyro rotors as a propulsion device. Cyclogyros are moreover, like helicopters, so-called vertical take-off aircrafts (also called VTOL vehicles, for “Vertical Take-Off and Landing”), i.e. aircrafts which are capable of starting and landing vertically without requiring a runway.
In stationary operation all blades of the cyclogyro rotor ideally are to be oriented best possible to the direction of flow at any time so as to make a maximum contribution to the entire thrust with minimally required driving power. The maximum pitch angle of the blades relative to the direction of flow has a direct influence on the amount of the thrust generated. Due to the rotation of the rotor the pitch angle of each blade has to be changed continuously during a revolution. Each blade of a cyclogyro rotor thus performs a periodic change of the pitch angle. This periodic change of the pitch angle is called pitch movement.
Various pitch mechanisms are known for pitch movement generation. For instance, each blade may be connected with an eccentric bearing axis via one or a plurality of conrods. The resulting pitch movement of a blade is repeated cyclically with every rotor revolution then. Consequently, the progression of the pitch angle may be developed into a Fourier series as a function of the current rotor twist. In this representation the fundamental harmonic value is typically dominating. It is superimposed by a mean value, and by higher harmonic values. The latter constitute undesired vibrations which stress the individual structural elements of a cyclogyro rotor. Since their amplitude and phasing cannot be chosen directly, they cannot be used for optimizing the aerodynamic efficiency, either.
Due to the quick rotation speed of the cyclogyro rotor in operation, its components are inter alia subject to loads in the form of forces of inertia and moments of inertia. This applies in particular for the blades since they are, as a matter of principle, very far from the axis of rotation, perform complex movements and form a relatively high share of the total mass of the cyclogyro rotor.
In the typical implementation of a cyclogyro rotor a part of the centrifugal force acting on the blade is introduced at one side of the conrod. The pitch movement forced by the conrod produces, owing to the mass inertia of the blade, additional forces in the first-mentioned. The second side of the conrod is connected with an eccentrically mounted offset disk (or directly with an offset pin).
Irrespective of the number of blades this results in load on the eccentric bearing axis in the form of a force acting radially outward. The time average of this force increases in good approximation linearly with the maximum pitch angle of the blade and by the square with the rotational speed.
This load constitutes a great challenge when designing a cyclogyro rotor. With respect to the installation space available and the lightweight construction requirements typical in aviation, it is not possible to design the eccentric bearing axis arbitrarily stable.
Typically, the position of the eccentric bearing axis is designed to be adjustable for changing the thrust. A necessary adjustment unit may be overloaded by the forces occurring. The consequence of this is that the eccentric bearing point moves further away from the axis of rotation, which results in a higher maximum pitch angle and consequently a higher load on the eccentric bearing axis. The result of this is an instable behavior which leads regularly to the destruction of the cyclogyro rotor. Additionally, the load on the eccentric bearing axis increases the energy consumption in the adjustment unit and restricts the dynamics thereof.
It is therefore an object of the present invention to reduce the afore-mentioned loads on the eccentric bearing axis of a cyclogyro rotor at high speed.
In accordance with a first aspect of the invention a propulsion device for an aircraft is provided which comprises the following components: a blade which can be rotated about an axis of rotation of the propulsion device along a circular path; a pitch mechanism having a coupling device and a bearing device. The blade is, by the bearing device, mounted for pivoting about a blade bearing axis parallel to the axis of rotation of the propulsion device. The propulsion device in accordance with the invention further comprises an offset device to which the blade is coupled by the coupling device at a connection point. The offset device defines an eccentric bearing axis which is mounted at an adjustable offset distance parallel to the axis of rotation of the propulsion device in such a way that the rotation of the blade about the axis of rotation of the propulsion device along the circular path effects a pitch movement of the blade when the offset distance is set to a nonzero value. The coupling device is coupled to the blade at a coupling point, wherein the coupling point is positioned in such a way that the plane that comprises the blade bearing axis and the coupling point and the tangential plane to the circular path through the blade bearing axis include a certain, non-vanishing angle when the offset distance is set to zero.
Due to the fact that the eccentric bearing axis which is defined by the offset device is mounted eccentrically at an offset distance parallel to the axis of rotation of the propulsion device, a pendular movement of the blade about the blade bearing axis of the blade results when the blade is coupled by the coupling device. This pendular movement is called pitch movement.
In accordance with the invention the pitch movement is described by the angle included by the tangent and/or tangential plane to the circular path through the blade bearing axis and the chord of the blade. It is of advantage if the pitch movement takes place in an angular range of −50° to +50° about the tangent to the circular path. When this angular range is used, relevant thrust forces can be generated. In the case of a symmetrical pitch movement with respect to the tangent to the circular path the blade is positioned in such a way relative to the blade bearing point and the coupling point that the chord and the tangent to the circular path are parallel when the offset distance is zero. If the chord is, with an offset distance zero, already positioned in a twisted manner relative to the tangent to the circular path, the result is a non-vanishing, but constant pitch angle with an offset distance zero, and consequently an asymmetrical pitch movement with respect to the tangent to the circular path with a non-vanishing offset distance. It may therefore be advantageous that the pitch movement takes place asymmetrically about the tangent to the circular path, i.e. in this case the maximum angle of the pitch movement above the tangent is larger than the maximum angle below the tangent, or vice versa.
Due to the fact that the certain, non-vanishing angle is set with an offset distance of zero, the definition of the certain, non-vanishing angle is unambiguous. In the case of a non-vanishing offset distance this angle would always change as a function of the pitch angle.
The chord means the connecting line between the leading edge and the trailing edge of a blade.
The leading edge and the trailing edge are given by the intersections of the camber line with the profile contour. The camber line (also referred to as skeleton line, curvature line or bending line) is a line consisting of the centers between the upper side and the lower side of the blade profile perpendicular to the chord. The camber line is in relation to the asymmetry between the upper side and the lower side of the blade profile. In the case of symmetrical profiles the camber line corresponds to the chord. Preferably, symmetrical profiles are used. The invention is not restricted to symmetrical profiles, though.
Due to the fact that the coupling point is positioned such that the plane that comprises the blade bearing axis and the coupling point and the tangential plane to the circular path through the blade bearing axis include a certain, non-vanishing angle when the offset distance is set to zero, higher harmonic values of the pitch movement of the blade can be influenced and reduced.
It is emphasized in this place that the effect in accordance with the invention occurs completely independently of the specific geometry and design of the blade and/or blade profile. In accordance with the invention, only the angle is important which is included by the tangential plane to the circular path through the blade bearing axis and the plane that comprises the blade bearing axis and the coupling point and/or which is included by the tangent to the circular path at the blade bearing point and the connection straight line through the blade bearing point and the coupling point.
A pitch movement occurs when the offset distance is set to a nonzero value. In accordance with the invention the coupling point is thus determined in a configuration in which the eccentric bearing axis which is defined by the offset device and the axis of rotation of the propulsion device are matching. For the operation of the propulsion device it is expedient to set the offset distance to a non-vanishing value to thus cause the pitch movement. When the non-vanishing offset distance exists, thrust is generated in a certain direction.
The pitch movement of the blade is repeated cyclically with every rotor revolution. The progression of the pitch angle may therefore be developed into a Fourier series as a function of the current rotor twist. In this representation the fundamental harmonic value is typically dominating. It is superimposed by a mean value, and by the afore-mentioned higher harmonic values.
The coupling point of the coupling device to the blade performs two rotational movements in the operation of the propulsion device. The first rotational movement occurs due to the rotation of the blade about the axis of rotation of the propulsion device. The second rotational movement is caused by the pitch mechanism which pivots the blade about the blade bearing axis. Due to the geometric construction of the pitch mechanism there result higher harmonic values in the second rotational movement, the pitch movement, and in continuation due to the superimposing with the first rotational movement higher harmonic values in the loads of the blade.
These higher harmonic values constitute unintended vibrations in the loads which may be transferred via the coupling device to the offset device and/or the eccentric bearing axis thereof. This impairs the stability of the offset device and of the eccentric bearing axis thereof.
Preferably, the coupling point is positioned at the side of the tangential plane to the circular path which faces the axis of rotation of the propulsion device.
It is of advantage if the coupling point is positioned to be shifted from the tangential plane of the circular path in the direction of the axis of rotation of the propulsion device in such a way that the certain, non-vanishing angle lies in a range of between 5° and 15°, preferably in a range of between 8° and 12°, particularly preferred in a range of between 9.5° and 10.5°.
Furthermore, it is of particular advantage if the certain, non-vanishing angle is set such that the plane that comprises the blade bearing axis and the coupling point and the plane that comprises the axis of rotation of the propulsion device and the connection line from the coupling point to the axis of rotation include an angle of almost 90° when the offset distance is set to zero. In this case all the even higher harmonic values of the pitch movement almost cancel out during the rotation of the propulsion device about its axis of rotation. Thus, the loads on the offset device are also minimized by the even higher harmonic values of the pitch movement.
It is expedient to determine the certain, non-vanishing angle as a function of the ratio of the two following dimensions: First: the distance of the blade bearing axis to the coupling point; second: the distance of the axis of rotation of the propulsion device to the blade bearing axis; each provided that the offset distance is set to zero. Preferably, the certain, non-vanishing angle indicated in radian assumes a value in the range of 75% to 125% of the afore-mentioned ratio of the first dimension to the second dimension; in a particularly preferred manner the certain, non-vanishing angle assumes a value in the range of 90% to 110% of the ratio mentioned.
Furthermore, it is of particular advantage if the certain non-vanishing angle is set such that the plane that comprises the blade bearing axis and the coupling point and the plane that comprises the axis of rotation of the propulsion device and the connection line from the coupling point to the axis of rotation include an angle of almost 90° when the offset distance is set to zero. In this case all the even higher harmonic values of the pitch movement almost cancel out during the rotation of the propulsion device about its axis of rotation. Thus, the loads on the offset device are also minimized by the even higher harmonic values of the pitch movement.
Preferably, the coupling point of the coupling device to the blade is positioned outside of the blade profile. This has the advantage that the blade as such is not impaired by the coupling of the coupling device. Advantageously the stability of the blade is thus not impaired adversely.
Preferably, blades are used with a profile which is symmetrical with respect to the chord. The invention is, however, not restricted to such symmetrical profiles.
It is of advantage if the blade bearing axis is shifted toward the axis of rotation of the propulsion device by a certain distance relative to the plane that extends through the center of mass of the blade and that extends parallel to both the axis of rotation and the chord of the blade.
The shifting of the blade bearing axis toward the axis of rotation by a certain distance relative to the plane that extends through the center of mass of the blade and that extends parallel to both the axis of rotation and the chord of the blade has the advantage that the mean force at the offset device and/or at the eccentric bearing axis which is defined by the offset device is minimized. The mean force at the offset device is the average of the entire force acting on the offset device in the course of a complete rotor revolution. If the force acting on the offset device and/or the eccentric bearing axis is developed into a Fourier series, the mean force is the term of zeroth order. Thus, the load on the offset device and/or on the eccentric bearing axis is further reduced. Loads are also exerted on the offset device and/or eccentric bearing axis due to torques engaging on the blade. As already explained before, the blade performs two rotational movements coupled with one another during the operation of the propulsion device. The first rotational movement originates from the rotation of the blade about the axis of rotation of the propulsion device along the circular path. The second rotational movement corresponds to the pitch movement of the blade about the blade bearing axis. In relation to the blade bearing axis two contributions result to the torque acting on the blade. The first contribution is related with the rotational movement of the blade about the axis of rotation of the propulsion device along the circular path. This rotational movement effects a centrifugal force on the blade. A corresponding torque is therefore always produced when the blade is mounted at a distance from its center of gravity. The second amount is related with the pitch movement of the blade about its blade bearing axis. The corresponding torque depends on the mass moment of inertia of the blade, on the one hand, and on the (angular) acceleration of the blade experienced during the pitch movement, on the other hand.
Both contributions to the torque mentioned are dependent on the distance of the blade bearing point and/or the blade bearing axis from the center of mass. The resulting torque may therefore be minimized by varying this distance.
The resulting torque produces diverse forces, such as tensile and/or compressive forces, in the coupling device. These forces are transferred to the offset device via the coupling device. By the positioning of the blade bearing point and/or the blade bearing axis in accordance with the invention it is thus possible to minimize the mean force at the offset device and/or the eccentric bearing axis.
Preferably, the blade has a mass distribution which is so inhomogeneous that it causes the shifting by the certain distance. A simple implementation of this inhomogeneous mass distribution consists in increasing the mass density on the blade upper side facing away from the axis of rotation, e.g. by applying additional weights or an appropriate coating on the blade upper side. Thus, the center of mass of the blade is displaced further outward in the radial direction relative to the axis of rotation of the rotor. With an otherwise unchanged blade geometry the effect according to the invention can thus be produced.
It is of advantage if the blade bearing axis is positioned in a region which is confined by the plane being perpendicular to the chord and extending through the center of mass, on the one hand, and by the plane being perpendicular to the chord and extending through the leading edge, on the other hand. This makes it possible to obtain exclusively tensile forces in the coupling device, which in turn enables a highly simplified construction thereof.
Preferably, the blade bearing axis extends outside the blade profile. This has the advantage that the stability of the blade is not impaired by the bearing device.
It is of advantage if the propulsion device in accordance with the invention further comprises a disk which designed such that it separates the blade(s) aerodynamically from the remaining components of the propulsion device. Such disk is especially advantageous for the case that the propulsion device is operated at higher speeds.
Preferably, the propulsion device further comprises a connection element, wherein the connection element, at the point at which the blade is mounted for pivoting by the bearing device, is connected rigidly with the blade, and at the coupling point of the blade is connected movably with the coupling device. The connection element comprises in a particularly preferred manner a lever arm. This enables the rigid connection of the lever arm with the blade. The connection element is preferably an independent structural element for coupling the coupling device to the blade. In a particularly preferred manner the lever arm is connected with the blade from the outside. The pitch movement is thus introduced into the blade via the bearing device. The advantage is that the coupling point of the coupling device may be chosen outside the blade profile. On the blade itself a place for mounting and for introduction of the pitch movement is sufficient. In a particularly preferred manner the blade bearing point is positioned at the place of the blade having the largest profile thickness. This has the advantage that the forces occurring can be distributed better in the blade. The result of this is also an improved construction and a weight reduction of the blade and/or of the pitch mechanism and/or of the propulsion device.
Specifically in the case of a propulsion device comprising a disk which is designed such that it separates the blade(s) aerodynamically from the remaining elements of the propulsion device, a connection element has the further advantage that no recess for the coupling device at the coupling point has to be provided in the disk. This is because in this case the coupling point can be chosen such that it does not get into contact with the disk.
It is advantageous if the offset device comprises an offset disk through the center of which the eccentric bearing axis extends, and which is mounted for rotating about the eccentric bearing axis, and wherein the connection point on the offset disk is arranged outside of the center thereof. In the case of a plurality of blades the offset disk comprises a corresponding connection point for each blade. The connection points are distributed evenly over the circumference of the offset disk. Instead of an offset disk it is also possible to use a so-called offset pin. In the case of a plurality of blades each blade is coupled to the same offset pin. The coupling devices of the individual blades are therefore coupled to the offset pin on top of each other. Thus, if an offset pin is used, the axial extension of the propulsion device increases as compared to the use of an offset disk.
In accordance with a second aspect of the invention a propulsion device for an aircraft is provided, comprising a blade which can be rotated about an axis of rotation of the propulsion device along a circular path; a pitch mechanism having a coupling device and a bearing device, wherein the blade is mounted by the bearing device for pivoting about a blade bearing axis parallel to the axis of rotation of the propulsion device; and an offset device to which the blade is coupled by the coupling device at a connection point. The offset device defines an eccentric bearing axis which is mounted at an adjustable offset distance parallel to the axis of rotation of the propulsion device, such that the coupling device couples the blade to the offset device in such a way that the rotation of the blade about the axis of rotation of the propulsion device along the circular path effects a pitch movement of the blade when the offset distance is set to a nonzero value. The blade bearing axis is shifted toward the axis of rotation of the propulsion device by a certain distance relative to the plane that extends through the center of mass of the blade and that extends parallel to both the axis of rotation and the chord of the blade.
The shifting of the blade bearing axis toward the axis of rotation of the propulsion device by a certain distance relative to the plane that extends through the center of mass of the blade and that extends parallel to both the axis of rotation and the chord of the blade has the advantage that the mean force at the offset device and the eccentric bearing axis is minimized Thus, the load of the offset device and of the eccentric bearing axis is further reduced. A detailed description of the loads exerted on the offset device due to forces engaging on the blade (the centrifugal force, on the one hand, inertia forces, on the other hand) has already been given further above in connection with an advantageous embodiment of the first aspect of the invention, which is referred to with respect to the second aspect of the invention. The two contributions mentioned are dependent on the distance of the blade bearing point and/or the blade bearing axis of the blade from the center of mass. By varying this distance it is thus possible to minimize the resulting force. This force is transferred to the offset device via the coupling device. Due to the positioning of the blade bearing point and/or the blade bearing axis at a particular distance from the center of mass it is thus possible to minimize the mean force at the offset device and/or the eccentric bearing axis.
Preferably, the blade has a mass distribution which is so inhomogeneous that it causes the shifting by the certain distance. A simple implementation of this inhomogeneous mass distribution consists in increasing the mass density on the blade upper side facing away from the axis of rotation, e.g. by applying additional weights or an appropriate coating on the blade upper side. Thus, the center of mass of the blade is displaced further outward in the radial direction relative to the axis of rotation of the propulsion device. This is equivalent to the blade bearing axis relative to the center of mass being closer in the radial direction to the axis of rotation of the propulsion device than the center of mass. With an otherwise unchanged blade geometry the effect according to the invention can thus be produced.
It is of advantage if the blade bearing axis is positioned in a region which is confined by the plane being perpendicular to the chord and extending through the center of mass, on the one hand, and by the plane being perpendicular to the chord and extending through the leading edge, on the other hand. This makes it possible to obtain exclusively tensile forces in the coupling device, which in turn enables a highly simplified construction thereof.
Preferably, the blade bearing axis extends outside the blade profile. Thus, the stability of the blade is not impaired by the bearing device.
Preferably, the coupling device comprises a conrod which connects the offset device with the coupling point of the blade. A conrod constitutes an implementation of the coupling device in accordance with the invention which is particularly suited in constructional respect. Preferably, an end piece of the conrod is coupled rotatably to the offset device.
It is of particular advantage to combine in one propulsion device advantageous embodiments of the first aspect of the invention with advantageous embodiments of the second aspect of the invention.
Preferably, the propulsion device in accordance with the invention comprises further blades, in a particularly preferred manner two, three, four, five, or six blades, with a respectively associated pitch mechanism, wherein all blades and pitch mechanisms of the propulsion device are of similar type, and wherein the blades of the propulsion device are evenly distributed about the axis of rotation of the propulsion device along the circular path. The propulsion device for an aircraft in accordance with the first or second aspect of the invention thus comprises preferably a plurality of blades which are distributed evenly about an axis of rotation of the propulsion device along a circular path, each of them being rotatable about the axis of rotation of the propulsion device along the circular path. Moreover, the preferred propulsion device comprises a plurality of pitch mechanisms with a respective coupling device and bearing device. Each blade is mounted by the corresponding bearing device for pivoting about a corresponding blade bearing axis parallel to the axis of rotation of the propulsion device. Furthermore, the preferred propulsion device comprises an offset device to which each blade is coupled by the corresponding coupling device at a corresponding connection point. The offset device defines an eccentric bearing axis mounted at an adjustable offset distance parallel to the axis of rotation of the propulsion device, so that the coupling devices couple the associated blades to the offset device in such a way that the rotation of the blades about the axis of rotation of the propulsion device along the circular path effects a pitch movement of the blades when the offset distance is set to a nonzero value. In accordance with the invention either each coupling device is coupled to the corresponding blade at a respective coupling point, wherein each of the coupling points is positioned such that the plane that comprises the corresponding blade bearing axis and the corresponding coupling point and the corresponding tangential plane to the circular path through the associated blade bearing axis include a certain, non-vanishing angle when the offset distance is set to zero. And/or the blade bearing axis of each blade is shifted toward the axis of rotation of the propulsion device by a certain distance relative to the plane that extends through the respective center of mass of the blade and that extends parallel to both the axis of rotation of the propulsion device and the chord of the respective blade.
In the case of a plurality of blades an even distribution about the axis of rotation of the propulsion device along the circular path means in connection with the invention that the blade bearing points and/or blade bearing axes of the blades are positioned approximately on the circular path and the blade bearing points and/or blade bearing axes of two adjacent blades each have almost the same distance from each other.
The use of a plurality of blades has the advantage that a higher thrust force of the propulsion device can be generated. Moreover, the even distribution of the blades along the circular path enables an at least partial cancelling of the forces acting on the offset device and/or the eccentric bearing axis. The advantageous embodiments of the first and/or second aspects of the invention may be applied correspondingly to the propulsion device with a plurality of blades. The resulting advantages correspond to those described in connection with the first and/or second aspects of the invention.
It is a particular advantage if the propulsion device comprises a total of five blades. Calculations have shown that propulsion devices in accordance with the invention which have a different number of blades react differently to harmonic values in the forces transferred by the corresponding coupling devices to the offset device and/or the eccentric bearing axis. Higher harmonic values occur, which load the offset device in the end. In the case of a total of five blades the higher harmonic values are strongly suppressed. This suppression is specifically intensified in a particular manner by the positioning of the coupling point of the coupling device at the blade in accordance with the invention, and/or by the positioning of the blade bearing axis in accordance with the invention.
Preferably the propulsion device is a cyclogyro rotor. The invention is not restricted to the use in cyclogyros, though. It is also possible to use the propulsion device in accordance with the invention e.g. in so-called Micro Air Vehicles (MAVs), i.e. unmanned drones of small size, or in manned aerial vehicles. Moreover, it is also possible to use the propulsion devices in accordance with the invention in connection with fluids other than air, such as, for instance, liquids.
In the following preferred embodiments of the present invention will be described by means of the following Figures. There show:
FIG. 1: a perspective view of an aircraft with a plurality of propulsion devices in accordance with the invention;
FIG. 2: a perspective view of a propulsion device in accordance with the invention;
FIG. 3: a profile view of a pitch mechanism coupled to an offset device in accordance with the first embodiment of the invention for defining the certain, non-vanishing angle;
FIG. 4: a profile view of a blade coupled to an offset device with a coupling device in accordance with the first embodiment of the invention;
FIG. 5a: a schematic diagram of the mode of functioning of the coupling of the coupling device at a coupling point of the blade in accordance with the invention;
FIG. 5b: a profile view of a blade coupled to an offset device with a coupling device, with a coupling point selected in an optimum position for reducing the loads on the offset device and/or eccentric bearing axis;
FIG. 5c: a propulsion device in accordance with the invention in a profile view at a rotor with four blades, wherein the coupling point is positioned optimally with all blades;
FIG. 6: a parameter study concerning the influence of the position of the coupling point of the blade on the loads on the offset device;
FIG. 7a: a first constructive variant of the first embodiment using a connection element for implementing the positioning of the coupling point in accordance with the invention;
FIG. 7b: a second constructive variant of the first embodiment using a connection to the blade for implementing the positioning of the coupling point in accordance with the invention;
FIG. 8: a profile view of a blade coupled to an offset device with a coupling device in accordance with the second aspect of the invention;
FIG. 9: a parameter study concerning the influence of the distance of the blade bearing point from the center of mass on the forces at the offset device;
FIG. 10: a Table indicating the loads on the offset device due to harmonic vibrations as a function of the number of blades.
FIG. 1 shows a perspective view of an aircraft 100 with a plurality of propulsion devices 1 in accordance with the invention. The illustrated aircraft 100 comprises four propulsion devices 1. The illustrated propulsion devices 1 are cyclogyro rotors. The aircraft 100 illustrated in FIG. 1 may therefore also be referred to as cyclogyro. The propulsion devices will be described in detail in connection with the following Figures. Each of these propulsion devices 1 is mounted for rotating about an axis of rotation. Each propulsion device 1 comprises a plurality of blades 2 which are mounted for pivoting about their longitudinal axes. Thus, it is possible to vary the pitch angle of the blades 2 during the rotation of the propulsion device 1. By controlling the speed of rotation of the propulsion devices 1 and the control of the pitch angles of the blades 2 it is possible to vary the amount and the direction of the thrust generated. The aircraft 100 comprises at its front side two major propulsion devices 1. At its rear side the aircraft 100 comprises to minor propulsion devices 1.
The illustrated aircraft 100 may, for instance, be an air vehicle, a manned aerial vehicle, a drone, or so-called Micro Air Vehicles (MAVs).
FIG. 2 illustrates a propulsion device 1 in accordance with the invention in a perspective view. This propulsion device 1 comprises five blades 2, respectively associated pitch mechanisms 3, an offset device 4 and a disk 11. The blades 2 are mounted for rotating about an axis of rotation of the propulsion device 1. The offset device 4 defines an eccentric bearing axis which is mounted eccentrically with respect to the axis of rotation of the propulsion device 1. In FIG. 2 the offset device is illustrated as an offset disk. The offset disc is mounted for rotating freely about the eccentric bearing axis. The eccentric bearing of the offset disk 4 implies an eccentric bearing of the pitch mechanism 3. The eccentric bearing of the pitch mechanism 3 effects the changing of the position of the blades 2 during a revolution about the axis of rotation of the propulsion device 1. Each of the illustrated pitch mechanisms 3 comprises a coupling device 31 and a bearing device 33. Each blade 2 is mounted for pivoting by the corresponding bearing device 33. The blade 2 is mounted about an axis parallel to the axis of rotation of the propulsion device 1. This axis is the blade bearing axis 33. The bearing of the blade 2 may, for instance, take place by a bearing means such as one or a plurality of pins, a so-called main pin. The bearing means is preferably a part of the bearing device 33. The blade bearing axis 33 may extend through the center of mass of the blade 2. Preferably, however, a bearing of the blade 2 takes place at a distance from the center of mass. The coupling device 31 of the pitch mechanism 3 couples the blade 2 to the offset device 4 in such a way that the blade 2 performs a pitch movement when it rotates about the axis of rotation of the propulsion device 1, and provided that the eccentric bearing axis does not coincide with the axis of rotation of the propulsion device 1. One end piece of the coupling device 31 is coupled to the offset device 4 at a connection point. The other end piece of the coupling device 31 is coupled to the blade 2.
The offset disk 4 is mounted for rotating freely. The axis of rotation of the offset disk 4 extends preferably at a certain offset distance parallel to the axis of rotation of the propulsion device 1. This produces the eccentric bearing of the offset disk 4 relative to the axis of rotation of the propulsion device 1. This offset distance may be adjustable. An offset device 4 with adjustable eccentricity may, for instance, be implemented by a planetary gear. A pitch movement of the blades 2 results when the offset distance is nonzero.
The coupling of the coupling device 31 to the blade 2 takes place at a coupling point 32. For this purpose the coupling device 31 may comprise a coupling means. In the propulsion device 1 illustrated in FIG. 2 the coupling device 31 comprises a conrod as well as a pin, the so-called pitch link pin. The pin is a constructive design of the coupling means in accordance with the invention. In the embodiment illustrated in FIG. 2 the coupling of the coupling device 31 to the blade 2 at the coupling point 32 is not performed by a direct connection with the blade 2, but by using a connection element 61. One end of the connection element 61 is rigidly connected with the blade 2. This connection takes place preferably at the blade bearing point. The other end of the connection element 61 is coupled to the coupling device/conrod 31. In this case the pitch movement is, via the coupling element by means of the conrod 31, introduced indirectly via the connection element 61 into the blade 2.
A direct coupling of the coupling device 31 to the blade 2 is, however, also possible in accordance with the invention.
Due to the fact that the coupling device 31 of the pitch mechanism is mounted eccentrically with respect to the axis of rotation of the propulsion device 1, the coupling point 32 moves relative to the blade bearing axis 33 on a circular arc when the blade 2 rotates about the axis of rotation of the propulsion device 1. This produces the pitch movement of the blade 2. It is thus a pendular movement of the blade 2 about the blade bearing axis 33.
Furthermore, the propulsion device 1 illustrated in FIG. 2 comprises a disk 11. This disk 11 is designed such that it separates the blades 2 aerodynamically from the remaining components of the propulsion device 1. Such a disk 11 is of particular advantage for the case that the propulsion device 1 is operated at higher speeds.
In an embodiment according to the invention in accordance with the first aspect of the invention the coupling device 31 is coupled to the blade 2 at a coupling point 32 which is positioned such that the plane that comprises the blade bearing axis 33 and the coupling point 32 and the tangential plane to the circular path through the blade bearing axis 33 include a certain, non-vanishing angle when the offset distance is set to zero. The blades 2 illustrated in FIG. 2 have a symmetric profile. A detailed description of the coupling device 31 in accordance with the invention will be found in particular in connection with FIG. 3.
The propulsion device 1 generates thrust due to two rotational movements coupled with one another. The first rotational movement is the rotation of the blades 2 about the axis of rotation of the propulsion device 1. This first rotational movement results in a movement of the blades 2 about the axis of rotation of the propulsion device along a circular path. Specifically, the blade bearing axes 33 and/or blade bearing points move along the circular path. Each blade bearing axis 33 is parallel to the longitudinal axis of the blades 2. The longitudinal axis of the blades 2 is parallel to the axis of rotation of the propulsion device 1. Thus, the longitudinal axis of the blades 2 is also parallel to the blade bearing axis 33. The direction of thrust of the propulsion device 1 is normal to the axis of rotation of the propulsion device 1. For an optimum thrust generation all blades 2 are to be oriented best possible to the direction of flow at any point of time. This ensures that each blade 2 makes a maximum contribution to the total thrust. During the rotation of the propulsion device 1 about its axis of rotation the pitch of each blade 2 is changed continuously due to the afore-described pitch mechanism. Each blade 2 performs a periodic change of the pitch angle and/or a pendular movement. This is the pitch movement. In this process the coupling point 32 moves on a circular arc about the blade bearing axis 33. This is the second rotational movement.
The amount and the direction of the thrust generated depend on the pitch of the blades 2. Therefore, the distance of the eccentric bearing of the offset device 4 and/or of the pitch mechanism 3 from the axis of rotation of the propulsion device 1 influences the amount of the thrust generated. By the shifting of the eccentric bearing of the offset device 4 in the circumferential direction, i.e. with a constant distance from the axis of rotation of the propulsion device 1, the direction of the thrust generated is changed.
Although in FIG. 2 pitch mechanisms 3 are illustrated at one side of the propulsion device 1 only, it may be expedient for reasons of stability to apply corresponding pitch mechanisms also at the opposite side of the propulsion device.
FIG. 3 shows a part of a propulsion device in accordance with the first aspect of the invention in a profile view. FIG. 3 illustrates a pitch mechanism and an offset device 4. The pitch mechanism comprises a coupling device 31 and a bearing device 33. Moreover, a part of the circular path 52 is indicated, along which the blade bearing axis 33 moves. Also illustrated is the tangent 54 to said circular path. The first aspect of the invention manages completely without the specific geometry of a blade. Therefore, no specific blade profile is drawn in FIG. 3. In accordance with the invention it is only the positioning of the coupling point 32 relative to the tangent 54 that is important. More specifically, the angle wα included by the tangent 54 and the connection straight line through the blade bearing point 33 and the coupling point 32 is crucial in accordance with the invention. The angle wα is determined in the configuration of the offset device in which no eccentricity exists, i.e. when the offset distance is set to zero, as illustrated in FIG. 3. This is equivalent with the axis of rotation 51 of the propulsion device coinciding with the eccentric bearing axis 41. The effect in accordance with the invention, i.e. the reduction of loads at the offset device 4 and/or at the eccentric bearing axis 41, occurs when the angle wα assumes a non-vanishing value. This will be shown below in connection with FIG. 6. Instead of the angle wα it may be expedient to consider the angle ρ which is included by the connection line of blade bearing axis 33 to coupling point 32, on the one hand, and by the connection line from the coupling point 32 to the axis of rotation 51, on the other hand.
It has to be noted that a so-called offset disk is shown in FIG. 3. The coupling device 31 is coupled at a connection point 42 of the offset disk which has a certain distance from the bearing axis 41 of the offset disk. Due to the fact that the offset disk is mounted for rotating, however, this does not yet result in a pitch movement. Only when the bearing axis 41 of the offset disk 4 is shifted relative to the axis of rotation 51 of the propulsion device radially by an offset distance does the pitch movement occur.
Finally, it has to be noted that FIG. 3 illustrates the part of the propulsion device in accordance with the invention as a cross-section in profile. When considering the extension of the propulsion device in the third, non-illustrated, dimension, planes will possibly have to be considered instead of straight lines when defining wα. In a general form there applies: The coupling point 32 is positioned such that the plane that comprises the blade bearing axis 33 and the coupling point 32 and the tangential plane 54 to the circular path 52 through the blade bearing axis 33 include a certain, non-vanishing angle wα when the offset distance is set to zero.
If, in the following, two-dimensional dimensions are referred to for simplifying matters, this implies that they are possibly mentioned representatively for corresponding three-dimensional dimensions.
FIG. 4 illustrates a part of a propulsion device in accordance with the first aspect of the invention in a profile view. FIG. 4 shows a blade 2, a pitch mechanism and an offset device 4. The pitch mechanism comprises a coupling device 31 and a bearing device 33. To simplify illustration, a blade 2 with a symmetrical profile will be considered here. The blade 2 is mounted for pivoting about the blade bearing axis 33 by means of the bearing device 33. The coupling device 31 is coupled to the blade 2 at the coupling point 32. The coupling point 32 is positioned such that the plane that comprises the blade bearing axis 33 and the coupling point 32 and the tangential plane 54 to the circular path through the blade bearing axis 33 include a certain, non-vanishing angle wα. The angle wα is determined in the configuration of the offset device in which no eccentricity exists, i.e. when the offset distance 43 is set to zero. The chord 230 is defined as the connection between the leading edge 210 and the trailing edge 220 of the blade. The definition of wα was already described above in connection with FIG. 3. The illustrated coupling device 31 is coupled with one end to the blade directly at the coupling point 32. That means that the coupling device 31 is, for instance, by using a coupling means such as a pin, directly connected with the blade 2 for moving. An indirect coupling of the coupling device 31 to the blade 2 is also possible; this will be described further below in connection with FIG. 7a.
The blade 2 may rotate along the axis of rotation 51 of the propulsion device about a circular path 52. The direction of rotation is indicated by the arrow 53; it is thus assumed that the blade rotates clockwise. The illustrated offset device 4 is an offset disk. It is mounted for rotating about an eccentric bearing axis 41. Preferably, the offset device 4 may rotate freely about this eccentric bearing axis 41. The eccentric bearing axis 41 of the offset device 4 is shifted parallel by a distance 43 relative to the axis of rotation 51 of the propulsion device. Due to this lateral displacement 43 the offset device 4 is mounted eccentrically relative to the axis of rotation 51. At the coupling point 42 of the offset disk 4 the coupling device 41 is coupled to the offset disk 4.
The chord 230 of the blade 2 illustrated in FIG. 4 is inclined by the angle α relative to the tangent 54 at the blade bearing point 33 to the circular path 52 which is described by the blade bearing point 33 during the rotation about the axis of rotation 51. This is the so-called pitch angle. The pitch movement occurs due to the eccentric bearing of the eccentric bearing axis 41 of the offset device 4 relative to the axis of rotation 51 of the propulsion device. FIG. 4 shows that, due to the eccentric bearing of the offset device 4, also the coupling point 42 of the coupling device 31 of the pitch mechanism passes along an eccentric path relative to the axis of rotation 51. The consequence of this is that the radial distance of the coupling point 42 changes during the rotation of the blade 2 relative to the circular path 52 on which the blade bearing axis 33 moves along. Thus, also the position of the coupling point 32 changes relative to this circular path 52. The effect of this is that the blade 2 performs a pitch movement α. In other words, the blade 2 lifts and lowers relative to the circular path 52. In even other words, the blade 2 performs a pendular movement about the circular path 52 while moving along the circular path 52. This pendular movement and/or the lifting and lowering of the blade 2 is indicated by the angle α. This is the so-called pitch angle. The angle α indicates the angle included by the tangent 54 at the blade bearing point 33 to the circular path 52 and the chord 230. It is of advantage to choose the maximum amplitude of the pitch movement such that the angle α can vary in a range of between −50° and +50°. Such angles are of advantage specifically for cyclogyro rotors so as to generate a relevant thrust.
The coupling point 32 moves during the pitch movement α on a circular arc about the blade bearing axis 33. This movement results in that the blade 2, during its movement along the circular path 52, performs a pendular movement about the axis of rotation 51 which comprises, in addition to a fundamental harmonic vibration, also higher harmonic values. These higher harmonic values are pronounced stronger the larger the pitch angle α becomes. In the case of the afore-mentioned angle range of −50° to +50° the higher harmonic values can no longer be neglected.
The positioning of the coupling point 32 in accordance with the invention provides a possibility of influencing the higher harmonic values mentioned.
FIG. 5a illustrates schematically how the positioning of the coupling point 32 in accordance with the invention leads to the influencing and reducing of the higher harmonic values of the pitch movement. In FIG. 5a the pitch mechanism is illustrated for the case that the axis of rotation 51 of the propulsion device matches with the eccentric bearing axis 41 of the offset device. In other words, the eccentricity of the eccentric bearing axis 41 relative to the axis of rotation is zero. In even other words, no pitch movement occurs in the illustrated case. In exactly this configuration the angle wα is defined and determined in accordance with the invention. A pitch movement introduced into the blade 2 by the coupling device 31 by coupling at the coupling point 32 only occurs when the eccentric bearing axis 41 of the offset device is again displaced by a certain distance from the axis of rotation 51 of the propulsion device. In this case the coupling point 32 moves along a circular arc 300 about the blade bearing axis 33.
In the case of a suitable choice of the angle wα the angle ρ which lies between the connection of the blade bearing point 33 with the coupling point 32, on the one hand, and the connection of the coupling point 32 with the axis of rotation 51 of the propulsion device and/or the eccentric bearing axis 41 of the offset device, on the other hand, is a right angle.
If the angle wα is chosen such that the angle ρ which is included by the connection line between the blade bearing point 33 and the coupling point 32, on the one hand, and the connection line between the coupling point 32 and the axis of rotation 51 of the propulsion device, on the other hand, is 90°, it results that the geometric deviation of the circular arc 300 from the tangent to the circular arc 300 at the coupling point 32 is distributed symmetrically. This means that, by the choice of a right angle, as described before, a symmetrization of the movement of the coupling point 32 relative to the tangent to the circular arc 300 is produced. This is equivalent with the fact that no even higher harmonic values exist in the pitch movement. All even higher harmonic values of the pitch movement can therefore be minimized by means of the angle wα.
The higher harmonic values of the pitch movement lead to forces in the coupling device 31. The coupling device 31 transfers these forces to the offset device and/or the eccentric bearing axis 41. This causes loads at the offset device. Due to the fact that the even higher harmonic values of the pitch movement are minimized by a positioning of the coupling point 32 in accordance with the invention, the corresponding loads at the offset device and/or at the eccentric bearing axis 41 are also minimized.
Although the optimum position of the coupling point 32 exists when the angle ρ which is included by the connection line between the blade bearing point 33 and the coupling point 32, on the one hand, and the connection line between the coupling point 32 and the axis of rotation 51 of the propulsion device, on the other hand, is 90°, a reduction of the loads at the offset device also occurs with other angles wα which do not necessarily result in a right angle ρ.
Corresponding model calculations will be described with respect to FIG. 6.
FIG. 5b illustrates a section from a propulsion device in accordance with the invention in profile, wherein the coupling point 32 is positioned optimally. This means that in the illustrated configuration the even higher harmonic values of the pitch movement are reduced best possible. FIG. 5b illustrates the general case with an asymmetrical blade profile. The angle ρ which is included by the connection line between the blade bearing point 33 and the coupling point 32, on the one hand, and the connection line between the coupling point 32 and the axis of rotation 51 of the propulsion device, on the other hand, is 90°. As described before in connection with FIG. 5a, the positioning of the coupling point 32 and hence the setting of the angle ρ takes place under the premise of vanishing eccentricity, i.e. the eccentric bearing axis 41 of the offset device 4 coincides with the axis of rotation 51 of the propulsion device.
Due to the finite dimensions of the offset device 4 the connection point 42 of the coupling device 31 to the offset device 41 does not coincide identically with the eccentric bearing axis 41 of the offset device 4. The radius r of the circular path 52 along which the blade 2 moves during the rotation about the axis of rotation 51 is regularly distinctly larger than the distance of the connection point 42 from the eccentric bearing axis 41. Therefore, it is also possible to optimize the angle ρ which is included by the connection line between the blade bearing point 33 and the coupling point 32, on the one hand, and the connection line between the coupling point 32 and the connection point 42 of the offset device 4, on the other hand, i.e. to set it to almost 90°, without suffering any loss in the reduction of loads at the offset device 4.
FIG. 5c shows a propulsion device in accordance with the invention in a profile view at a rotor with four blades 2. Each of the blades 2 may move about the axis of rotation 51 of the propulsion device on a circular path 52 with the radius r. Each blade 2 is coupled to the offset device 4 at a coupling point 42. In the illustrated example the blades 2 rotate clockwise, as indicated by the arrow 53. In FIG. 5c the propulsion device is shown with non-vanishing eccentricity, i.e. the eccentric bearing axis 41 is shifted from the axis of rotation 51 by a non-vanishing offset distance 43. As explained already further above, this results in a pitch movement α of the blades 2. The coupling point 32 of each blade 2 has been positioned optimally. This means, as described especially with respect to FIGS. 3, 5a and 5b, the angle included by the connection straight line through the blade bearing point 33 and the coupling point 32 as well as by the coupling device 31 was set to almost 90. This angle was determined in the configuration of the offset device 4 in which no eccentricity exists, i.e. when the offset distance was set to zero.
Although the specific geometry, bearing, or the specific profile of the blades 2 is not important for achieving the effect in accordance with the invention—as already explained above—(crucial is the relative arrangement of blade bearing point 33, coupling point 32 and coupling device 31 with vanishing offset distance 43), it is assumed in the embodiment illustrated in FIG. 5c that the chords of the blades 2, with an offset distance zero, did not show any twist relative to the tangent 54 to the circular path 52. With this initial configuration the occurrence of the minimization of the even harmonic values of the pitch movement can be illustrated particularly well. This is because with the illustrated rotor with four blades 2 with optimal positioning of the coupling point 32 the minimization of the even harmonic values of the pitch movement manifests itself such that the opposing blades 2 each have the negative pitch angle α and/or −α of the other one. In the illustrated position the uppermost blade is at its maximum positive deflection a, the lowermost blade at its maximum negative deflection −α. The two other blades are in a middle position with a deflection of zero degrees.
For the sake of completeness it is mentioned that, due to the fact that the even higher harmonic values, under real conditions of operation, are only minimized and do not disappear completely, the opposing blade only comprises approximately the negative deflection value of the respectively other one.
FIG. 6 shows a graph 7 indicating the peak-to-peak value of the load at the offset device, Avib,hub, normalized to the peak-to-peak value of the loads at the offset device for wα=0, Avib,hub,0 a function of the angle wα. The peak-to-peak value describes the difference between the minimum and maximum values of the load at the offset device and is thus a direct measure for the vibration of the load at the offset device. The ordinate 71 indicates the function value Avib,hub/Avib,hub,0, the abscissa 72 the angle wα measured in degrees.
The loads at the offset device and/or at the eccentric bearing axis as a function of the value wα were calculated by using a further calculation of all forces and moments and an additional consideration of aerodynamic loads. The reduction of the loads at the offset device can be recognized clearly. Specifically, it results from the progression of the graph 7 that a reduction of the loads at the offset device occurs as soon as the coupling point is positioned in accordance with the invention. In operation, the angle between a coupling device and the connection line between the coupling point and the blade bearing point must be sufficiently acute at any point of time. Otherwise, self-retention would occur and the function of the pitch mechanism would no longer be given. Experience has shown that, with respect to a maximum pitch angle of 50°, the twist wα is restricted to maximally 20°. This means that an improvement compared to the coupling of the coupling device at the tangent to the circular path (considered for the case that no eccentricity exists) always occurs under realistic conditions of use in accordance with the invention.
FIG. 6 illustrates that the minimum of the load at the offset device occurs at an angle wα of approximately 10°. Taking into account the geometry underlying the model calculation, this corresponds actually approximately to a right angle between the connection line of blade bearing point to coupling point, on the one hand, and the connection line of coupling point to the axis of rotation of the propulsion device, on the other hand.
FIGS. 7a and 7b illustrate two variants in accordance with the invention for coupling the coupling device 31 to the blade 2. In both FIGS. 7a and 7b symmetrical blade profiles are shown. It is assumed that the offset distance is set to zero.
FIG. 7a illustrates an indirect coupling of the coupling device 31 to the blade 2. This means that the coupling point 32 is not positioned directly at the blade 2. The illustrated coupling point 32 is positioned outside of the blade profile. Coupling takes place via a connection element 61. The connection element 61 may be a lever arm. The connection element is connected with one end rigidly to the blade 2. It is illustrated that the connection element 61 is connected with the blade 2 with the aid of the bearing device 33, preferably with a bearing means such as a pin, the so-called main pin. The other end of the connection element 61 is movably connected with the coupling device 31 at the coupling point 32. The coupling point 32 is positioned away from the tangent 54 by an angle wα.
Due to the use of the connection element 61 as a separate structural element for coupling the coupling device 31 at the blade 2 only one bearing means, such as for instance a main pin, is required. This means that the stability of the blade 2 is impaired by the bearing device 33 at one place only, a second load such as for fastening the coupling device 31 directly at the blade 2 by an appropriate coupling means, such as for instance a further pin, is thus omitted. Due to the fact that the connection element 61 is connected rigidly at the outside with the bearing device 33 or with the bearing means and hence with the blade 2, the pitch movement is introduced into the blade 2 via a moment in the bearing device 33 and/or in the bearing means.
The variant shown in FIG. 7a yields several advantages. First, the optimum positioning of the coupling point 32 of the coupling device 31 at the blade 2 can be implemented in a particularly simple manner by the connection means 61. The angle can simply be adjusted by twisting the connection element 61 about the blade bearing axis 33 such that the angle ρ which is included by the connection element 61 and the coupling device/conrod 31 is approximately 90°. As described with respect to FIGS. 5a and 5b, an almost optimum position of the coupling point 32 is thus determined.
Furthermore, the total weight of the pitch mechanism together with the connection means 61 is lower than with a conventional direct coupling of the coupling device 31 to the blade 2 since the additional coupling means, such as an additional pin, for the introduction of forces is omitted. Moreover, the introduction of forces for the pitch movement takes place via the bearing device 33 and/or via the bearing means and hence regularly at the thickest place of the blade 2. Thus, the forces occurring can be better distributed in the blade 2. This in turn enables an improved construction and a further weight reduction of the propulsion device.
Finally, a further advantage results with propulsion devices providing a disk for the aerodynamic separation of the blades from the rest of the components of the propulsion device, such as it is for instance illustrated in FIG. 2 (designated with reference number 11 there). The coupling of the coupling device 31 to the blade 2 with the aid of a connection element 61 avoids the necessity of providing an additional recess in the disk for the connection of the coupling device 31 at the blade 2. Thus, a simpler construction of the disk is enabled. Moreover, the aerodynamics of the propulsion device is improved.
FIG. 7b shows a further variant of a connection element 62 for the coupling of the coupling device 31 at a coupling point 32 to the blade 2. The coupling point 32 is spaced apart from the tangent 54 by the angle wα. The coupling point 32 is positioned outside of the blade profile. The connection element 62 is fastened at the lower side of the blade at a position far from the bearing device 33. One end of the coupling device 33 is mounted for moving at the coupling point 32 of the connection element.
FIG. 8 shows a blade 2, a pitch mechanism and an offset device 4 of a propulsion device according to the second aspect of the invention. The profile of the blade 2 illustrated in FIG. 8 is asymmetrical. The part of a propulsion device in accordance with the invention illustrated in FIG. 8 differs from the corresponding part of the propulsion device illustrated in FIG. 4 in that the blade bearing axis 33 is arranged at a certain distance w from the center of mass 250 of the blade 2. More specifically: The blade bearing axis 33 is, relative to the plane 260 extending through the center of mass 250 of the blade and parallel to both the axis of rotation 51 and the chord 230 of the blade, shifted by the distance w toward the axis of rotation 51 of the propulsion device. In FIG. 8 the propulsion device is shown with an eccentric bearing axis 41 which is shifted from the axis of rotation 51 by an offset distance 43.
The chord 230 is defined as the connection line between the leading edge 210 and the trailing edge 220 of the blade 2. The leading edge 210 and the trailing edge 220 are given by the intersections of the camber line 240 with the profile contour. The camber line 240 is in turn defined as the line consisting of the centers between the upper side 241 and the lower side 242 of the blade profile perpendicular to the chord 230.
The generation of the pitch movement α by means of the pitch mechanism which comprises a coupling device 31 and a bearing device 33 takes place as described in connection with FIG. 4. The blade bearing axis 33 rotates along the circular path 52 at a distance r from the axis of rotation 51 of the propulsion device. The repeated description of the movement of the coupling point 32 of the coupling device 31 at the blade 2 and of the coupling point 42 of the coupling device 31 at the offset device 4 during the rotation of the propulsion device in the direction of the arrow 53 is therefore spared. Everything that was said before in connection with the pitch mechanism is also true for the embodiment according to the second aspect of the invention illustrated in FIG. 8.
In FIG. 8 the blade bearing axis 33 is shifted toward the axis of rotation 51 by a distance w from a straight line (and/or from the corresponding plane 260, if the extension of the propulsion device in the third dimension is taken into account) which extends through the center of mass 250 of the blade and parallel to the chord 230 of the blade. It is to be understood that the relevant dimensions were here considered with respect to their projection on a plane perpendicular to the axis of rotation 51. When making the three-dimensional extension of the propulsion device a basis, there applies in accordance with the invention: The blade bearing axis 33 is shifted toward the axis of rotation 51 by a particular distance wgx relative to the plane 260 which extends through the center of mass 250 of the blade and parallel to both the axis of rotation 51 and the chord 230 of the blade 2. Moreover, the blade bearing axis 33 is shifted by the distance wgz relative to the plane which is perpendicular to the chord 230 and which extends through the center of mass 250. It turns out that the distance wgz substantially influences the mean value of the loads in the coupling device 31. As for the rest, wgz has a negligible influence on the loads at the offset device 4.
The shifting wgx of the blade bearing axis 33 away from the center of mass 250 enables to reduce the first harmonic vibration of the torque at the blade 2. This will be explained in detail soon. The reduction of the first harmonic vibration is associated with a reduction of the mean force at the offset device 4. This will be explained in detail in connection with FIG. 10.
In the following, the influence of the distance wgx on the load at the offset device 4 will be described. The blade 2 is mounted for pivoting about the blade bearing axis 33 and/or at the blade bearing point 33. During the rotation of the propulsion device about the axis of rotation 51 the blade 2 performs two rotational movements. The first rotational movement is the rotation of the blade 2 along the circular path 52, the second is the rotation of the blade 2 about the blade bearing axis 33 due to the pitch movement a. Each of these rotational movements effects a corresponding force and/or a corresponding torque on the blade 2. Due to the rotation of the blade 2 about the axis of rotation 51 of the propulsion device the centrifugal force FZ acts on the blade 2. This centrifugal force FZ engages in the center of mass 250 of the blade 2. If M designates the mass of the blade 2, r indicates the distance of the blade bearing point 33 from the axis of rotation 51 and ω indicates the angular speed of the propulsion device, then the amount of the centrifugal force FZ is given by
FZ=M·r·ω<2>.
The centrifugal force FZ in turn effects a torque TZ at the blade 2 which attempts to rotate the blade 2 about the blade bearing axis 33. This torque TZ is given by
TZ=FZ·
Image available on "Original document"
,
wherein
Image available on "Original document"
is the distance of the center of mass 250 from the blade bearing axis 33; this means that
Image available on "Original document"
is given as the perpendicular from the blade bearing axis 33 on the vector of the centrifugal force FZ engaging in the center of mass 250. The distance
Image available on "Original document"
depends on the pitch angle α of the blade; in other words, the distance Image available on "Original document"
is a function of the pitch angle α; in even other words, the distance Image available on "Original document"
is a function of the pitch movement α. Therefore: Image available on "Original document"
=Image available on "Original document"
(α).
In addition to the torque TZ caused by the centrifugal force FZ, another torque TI acts on the blade 2 due to the pitch movement α about the blade bearing axis 33. This torque TI depends, on the one hand, on the mass inertia moment I of the blade, relating to the blade bearing axis 33, and, on the other hand, on the angular acceleration of the pitch movement α. The torque TI is given as
[mathematical formula]
wherein the angular acceleration is given by the second time derivative of the pitch movement α.
The total torque T acting on the blade is thus given by
[mathematical formula]
A Taylor expansion of the total torque T in the pitch angle α and/or in the pitch movement α results in that, with realistic amplitudes αA of the pitch movement α, such as for instance αA=50°, i.e. −50°<α<+50°, harmonic values of the pitch movement which are higher than the fundamental harmonic vibration can substantially be neglected with respect to the mean force on the eccentric bearing axis 41. Moreover, it results from the Taylor expansion in consideration of the geometry illustrated in FIG. 8 that the contribution to the total torque T on the blade 2 produced by the fundamental harmonic vibration is proportional to the following term R:
[mathematical formula]
Icm designates the mass inertia moment calculated with respect to the center of mass 250 of the blade which may be calculated by means of the Steiner theorem from the mass inertia moment I, related to the blade bearing axis 33. wgz indicates the distance of the center of mass 250 from a straight line which is perpendicular to the chord 230 and extends through the blade bearing point 33. It turns out that wgz substantially influences the mean value of the moment. Thus, wgz can be used to influence the mean values in the coupling device 31. This coupling device 31 is, due to geometry, regularly a very large structural element in which a compressive load may cause failure by kinking, which thus constitutes a critical loading condition. With the parameter wgz a bias can now be effected in the tension direction in the coupling device 31 by the shifting of the mean values, so that no compressive forces occur therein in operation. Thus, the critical loading condition of a compressive load need not be taken into consideration, and this structural element can be designed in a substantially simpler way. wgz has moreover no substantial influence on the load of the eccentric bearing axis 41. Due to the symmetrical distribution of a plurality of coupling devices 31 along the circular path 52 the mean values in the coupling devices 31 in the offset device 4 cancel out. In a particularly preferred manner wgz is chosen such that the blade bearing point lies between the center of mass 250 and the leading edge 210. wgx influences substantially the first harmonic value of the torque at the blade. For the determination of this first harmonic value wgz may be neglected. This is due to the fact that wgz is indeed contained in the formula for optimization (term R above), but that it has only a very small influence as compared to wgx. This can be seen by the fact that wgz is only included by the square in the above formula for R.
If this term R is minimized, i.e. R=0, the torque T of the blade is also minimized. The influence exerted by the torque T on the blade is transferred via the coupling device 31 to the offset device 4. In connection with FIG. 10 it will be described that the first harmonic value of the torque T at the blade 2 results in a mean force at the offset device 4. This means that it is possible by the variation of the distance wgx to minimize the loads at the offset device 4 due to the centrifugal force FZ and the mass inertia of the blade 2.
FIG. 9 illustrates the progression of the two components 91, 92 of the mean force at the eccentric bearing axis in x and/or z direction (global coordinate system), both normalized to the force Fx at wgx=0. Thus, also the relation between the components may be read. The ordinate 93 indicates the function value, the abscissa 94 the distance wgx measured in centimeters.
In this parameter study the eccentric bearing axis is deflected in the positive x direction. Thus, a substantial mean force results at wgx=0 in the x direction. This may now be reduced with the parameter wgx. In the case of an ideal configuration this component may even vanish. Preferably wgx is chosen such that the component Fx 91 becomes negative. Thus, a stabilization of the system results since the mean force at the eccentric bearing axis counteracts the deflection thereof. If the offset distance is now increased, the mean force also increases, which precisely counteracts this deflection. However, at wgx=0 the component Fx 91 acts in the direction of the deflection. If the deflection is increased, the mean force again also increases in the direction of the deflection, which corresponds to an instable property. If, for instance, during a failure of the control the eccentric bearing axis could move freely, the rotor would destroy itself at wgx=0 since the mean force always acts in the positive deflection direction. If, however, the force is directed contrary to the deflection, this has a stabilizing effect. The reduction of the mean force at the offset device can be recognized clearly. Specifically, the progression of the graph Fx reveals that a reduction of the mean force occurs at the offset device as soon as the blade bearing axis and/or the blade bearing point is positioned at a distance w from the center of mass of the blade. This means that an improvement as compared to the bearing of the blade in the center of mass always occurs in accordance with the invention when realistic conditions are made a basis.
The mean force at the offset device as a function of the distance w was calculated by using a further calculation of all forces and moments and an additional consideration of aerodynamic loads.
FIG. 9 reveals that the zero-crossing of Fx 91 of the mean force at the offset device occurs at a distance wgx of approximately 3.4 mm. Taking into account the geometry underlying the model calculation, this corresponds very well to the value obtained by minimizing the term R derived in connection with FIG. 8.
The embodiments described in particular in connection with FIGS. 4-6 and concerning the first aspect of the invention allow the minimization of the contributions of the even higher harmonic values of the vibrations of the torque at the blade. Due to the overlapping when a plurality of blades are used, for instance with five blades, a minimization of the total vibration thus results at the offset device, as will be shown in detail in connection with FIG. 10. The embodiments described in connection with FIGS. 8 and 9 and concerning the second aspect of the invention allow the minimization of the fundamental vibration of the torque at the blade, and further due to the overlapping of a plurality of blades, of the mean force at the offset device and/or the eccentric bearing axis.
By combining the first and second aspects of the invention it is therefore possible to substantially reduce the loads at the offset device of the propulsion device. This means that, by an appropriate choice of the angle wα and of the distance wgx, a substantial reduction of the vibrations and of the mean force at the offset device and/or the eccentric bearing axis and of the related loads is achieved.
FIG. 10 shows a Table demonstrating the influence of harmonic values on the load at the offset device and/or the eccentric bearing axis as a function of the number of blades of the propulsion device. The parameter n 81 designates the number of blades. The parameter j 83 indicates the ordinal number of the harmonic values of the loads at the offset device resulting from individual blades, wherein the loads were calculated in a reference system co-rotating with the propulsion device. If one proceeds to a stationary reference system, a redistribution and an overlapping of the harmonic values j 83 of all blades will result therefrom. In the stationary reference system the parameter k 82 designates the ordinal number of the harmonic values of the loads at the offset device. The Table indicates for each harmonic value k 82 in the stationary reference system which harmonic values j 83 in the co-rotating reference system determine same.
For the stationary reference system the following can be derived from the Table of FIG. 10. Irrespective of the number of blades a fundamental harmonic value in the co-rotating reference system always results in a mean force at the offset device. This becomes clear by the entries in the column designated with 84. In other words, the fundamental harmonic values j=1 in the load of the blade in the co-rotating reference system effect a mean force, characterized by the contribution of zeroth order k=0, in the stationary reference system. Furthermore, propulsion devices with different numbers of blades react differently to harmonic values j 83 in the loads in the co-rotating reference system. This results already from the fact that in the stationary reference system different harmonic values k vanish; vanishing harmonic values are designated by empty fields 87.
The Table of FIG. 10 finally reveals that a propulsion device with n=5 blades is particularly advantageous. This is first of all due to the fact that for the case of n=5 the harmonic values of the loads with the ordinal number k=1, 2, 3, 4 vanish in the stationary reference system. The harmonic values 86 with the high ordinal numbers k=10 and k=15 are strongly suppressed. Pursuant to the Table of FIG. 10 the loads at the offset device therefore result in the case of n=5 blades substantially from the mean force k=0, 84, and the harmonic value of fifth order, k=5, 85. The Table further reveals that the mean force 84 in the stationary reference system is effected by the fundamental harmonic vibration, j=1, in the blade and/or in the coupling device in the co-rotating reference system. This mean force can, as described before in connection with FIGS. 8, 9, be minimized by a positioning of the distance w at a particular distance from the center of mass of the blade. The harmonic value of fifth order results pursuant to FIG. 10 from the harmonic values of the fourth, j=4, and sixth, j=6, orders of the vibrations in the co-rotating reference system. These are even higher harmonic values. As described in connection with the embodiments illustrated in FIGS. 3-7, these values may be minimized by a choice—in accordance with the invention—of the coupling point of the coupling device at the blade at a particular angle from the tangential plane through the blade bearing point.
This shows that the two aspects in accordance with the invention effect a particularly advantageous reduction of the loads at the offset device and/or at the eccentric bearing axis with a propulsion device comprising five blades.
LIST OF REFERENCE NUMBERS
1 propulsion device
100 aircraft/cyclogyro
11 disk of the propulsion device 1
2 blade
210 leading edge of the blade 2
220 trailing edge of the blade 2
230 chord of the blade 2
240 camber line of the blade 2
241 upper side of the blade 2
242 lower side of the blade 2
250 center of mass of the blade 2
260 plane passing through the center of mass 250 and extending parallel to the axis of rotation 51 and parallel to the chord 230
3 pitch mechanism
31 coupling device of the pitch mechanism 3/conrod
32 coupling point of the coupling device 31 to the blade 2
33 bearing device of the pitch mechanism 3/blade bearing axis/blade bearing point
300 circular arc of the pitch movement
4 offset device
41 eccentric bearing axis
42 coupling point of the coupling device 31 to the offset device 4
43 offset distance of the eccentric bearing axis 41 from the axis of rotation 51 of the propulsion device 1
51 axis of rotation of the propulsion device 1
52 circular path about the axis of rotation 51
53 arrow for indicating the direction of rotation of the propulsion device 1
54 tangential plane/tangent to the circular path 52 through the blade bearing axis 33
61 connection element for the indirect coupling of the coupling device 31 to the blade 2
62 connection element for the coupling of the coupling device 31 to the blade 2
7 graph of the normalized loads at the offset device 4 as a function of the angle wα
71 ordinate, designating the normalized loads at the offset device 4
72 abscissa, designating the angle wα in degrees
81 number n of blades
82 ordinal number k of the harmonic values of the loads at the offset device indicated in a stationary reference system
83 ordinal number j of the harmonic values of the loads at the offset device indicated in a reference system co-rotating with the propulsion device
84, 85, 86 non-vanishing contributions to the load at the offset device
87 vanishing contributions
9 graph of the normalized mean force at the offset device 4 as a function of the distance wgx of the blade bearing axis 33 from the center of mass 250
91 x component of the mean force at the eccentric bearing axis 41
92 y component of the mean force at the eccentric bearing axis 41
93 ordinate, designating the normalized mean force at the offset device 4
94 abscissa, designating the distance wgx in millimeters
α pitch angle/pitch movement
wα angle between the tangent 54 to the circular path and the connection line of the coupling point 32 to the blade bearing axis 33
wgx distance of the blade bearing axis 33 from the plane 260 through the center of mass 250 and parallel to the chord 230
wgz distance of the blade bearing axis 33 from the plane through the center of mass 250 and perpendicular to the chord 230
r distance of the blade bearing axis 33 from the axis of rotation 51 of the propulsion device 1
Image available on "Original document"
distance of the center of mass 250 from the blade bearing point 33
ρ angle between blade bearing axis 33, coupling point 32 and axis of rotation 51
FZ centrifugal force acting on the blade
US10822079 -- AIRCRAFT [ PDF ]
Inventor: HOFREITHER KLEMENS [AT] // KINAST LUKAS [AT]
Applicant: CYCLOTECH GMBH
The invention relates to an aircraft designed as a compound helicopter with an aircraft fuselage (1), a main rotor (2) arranged on the aircraft fuselage (1), and cylcogyro rotors (3, 3') which protrude laterally from the aircraft fuselage (1) and which comprise an outer end surface. An improved torque compensation is achieved in that the cyclogyro rotors (3, 3') are connected to the aircraft fuselage (1) by means of a suspension device (4, 4') which holds the cyclogyro rotors (3, 3') at the outer border of the rotors, and each cyclogyro rotor (3, 3') can be controlled individually and independently of the other. A torque compensation function of the main rotor (2) can be carried out by the cyclogyro rotors (3, 3').[AT]
BACKGROUND OF THE INVENTION
Field of the Invention
[0002] The invention relates to an aircraft designed as a compound helicopter with an aircraft fuselage, a main rotor arranged on the aircraft fuselage, and cyclogyro rotors which protrude laterally from the aircraft fuselage and comprise an outer end surface.
Background Information
[0003] Cyclogyro rotors is, in general, the denotation for cylindrical bodies which are mounted in such a manner that they are rotatable about their axis and at the circumference of which there are arranged pivotable rotor blades which are cyclically adjusted by means of an offset adjusting device during operation. Thus a thrust can be generated in any direction perpendicular to the axis in dependence on the adjustment of the rotor blades.
[0004] Prior art is represented by compound (hybrid) helicopters which consist of an aircraft fuselage, a single main rotor or a counter-rotating tandem motor, one or several propeller units for the torque compensation and for the thrust generation in forward flight, as well as of additional wing or airfoil units for the generation of a vertical lift in forward flight. Furthermore, there are known helicopter configurations which comprise one or two cyclogyro rotors.
[0005] In the lateral arrangement of two rotors below the main rotor on the left-hand side and the right-hand side of the helicopter fuselage, respectively—as is known from prior art—the cyclogyro rotors are connected to the helicopter fuselage exclusively via the rotor shaft. As a result thereof, high forces and torques or moments will occur at the mounting on the aircraft fuselage and in the rotor shaft. Moreover, the cyclic adjustment of the rotor blades by means of a unilateral offset adjustment device is problematic, as enormous centrifugal forces will be generated due to the required high rotor speeds, and as additional torsional moments will burden the rotor blade disproportionally due to the unilaterally initiated cyclic rotor blade articulation.
[0006] In the following, further known solutions in connection with the torque compensation in helicopters will be discussed.
[0007] From EP 2 690 011 A (Axel Fink) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with two wings at each of which a thrust propeller is arranged backwardly and rigidly in the flight direction. The wings or airfoils are rigidly connected to the aircraft fuselage by means of struts. Instead of a tail rotor there is provided a horizontal and vertical stabilizer. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight. A similar aircraft configuration is known from U.S. Pat. No. 3,385,537 A.
[0008] From EP 2 690 010 A (Axel Fink) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with two wings which are connected towards the rear to the horizontal and vertical stabilizer by means of a double fuselage, wherein at the rear ends of the double fuselages a thrust propeller is rigidly arranged, respectively. The wings are rigidly connected to the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight.
[0009] From EP 2 690 012 A (Axel Fink) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with four wings, wherein at both front ends of each of the wings there is arranged a pivotally designed ducted fan, respectively. The wings are rigidly connected to the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated by the main rotor and is supported by the two ducted fans which also generate the torque compensation and the thrust in forward flight. The rear wings are provided with elevators and rudders, the front wings are provided with ailerons.
[0010] From EP 2 666 718 A (Paul Eglin) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with two wings and a horizontal stabilizer, wherein at the front ends of the wings propellers are rigidly arranged in the flight direction. The wings are rigidly connected to the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight.
[0011] From EP 2 146 895 A (Philippe Roesch) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with two wings and a horizontal and vertical stabilizer, wherein at the front ends of the wings propellers are arranged rigidly in the flight direction. The wings are rigidly connected with the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight.
[0012] From EP 2 105 378 A (Jean-Jaques Ferrier) there is known an aircraft configuration which is designed with an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, and with four wings wherein backwards at the larger rear wings there is rigidly arranged a thrust propeller in the flight direction, respectively. The wings are rigidly connected to the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight. The wings are additionally provided with elevators.
[0013] From DE 10 2012 002 256 A (Felix Fechner) there is known an aircraft which is designed as a helicopter with additional wings, wherein said wings are implemented to be pivotable or are implemented in segments and thereby produce a reduction of the obstruction of the rotor downwind and facilitate a higher flying velocity during hover flight or low speed flight. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor.
[0014] From RU 2 500 578 A (Nikolaevich Pavlov Sergej) there is known an aircraft configuration which is designed with a an aircraft fuselage at which a main rotor is provided approximately in the mass center of gravity, with two propeller units which are arranged in the front region laterally in relation to the aircraft fuselage and in parallel to the flight direction for the forward thrust and with two pivotable wings as horizontal stabilizer, and a vertical stabilizer in the rear region. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while the two additional propellers generate the torque compensation and the thrust in forward flight.
[0015] From US 2013 0327879 A (Mark W. Scott) there is known an aircraft configuration which is designed as a helicopter with a main rotor and a tail rotor, which can be pivoted about an axis of rotation, approximately in parallel to the axis of rotation of the main rotor. The pivotable tail rotor stabilizes the aircraft in the hovering state and it can additionally generate a horizontal thrust in the flight direction, while during take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor.
[0016] From US 2006 0169834 A (Allen A. Arata) there is known an aircraft configuration which is designed as a helicopter with a main rotor and with a tail rotor, and with two additional wings. The wings are rigidly arranged at the aircraft fuselage below the main rotor and can be pivoted approximately in the middle of their length by 90° downwards in parallel to the aircraft axis, and in this position they serve as landing skid or landing gear. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two extended wings.
[0017] From WO 2005/005250 A (Arthus W. Loper) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the front end of the helicopter, with two additional wings and with a horizontal and vertical stabilizer. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The front-end propeller generates the thrust for the forward flight.
[0018] From US 2006 0157614 A (John S. Pratt) there is known an aircraft which is designed as a helicopter with several additional wings below the main rotor, wherein said wings are implemented in segments and in a pivotable manner, and thereby they enable a reduction of the obstruction of the rotor downwind and facilitate a higher flying velocity during hover flight or low speed flight. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, and in the faster forward flight the additional wings support the vertical lift. The torque compensation is carried out by means of the individual setting of the segmented wings via the downwind of the main rotor, and no tail rotor is present.
[0019] From FR 9 803 946 A (Paul Julien Alphonse) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, with two additional wings and with a horizontal and vertical stabilizer. The wings are rigidly arranged at the aircraft fuselage outside the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight.
[0020] From U.S. Pat. No. 5,738,301 A (Daniel Claude Francois) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, with two additional wings, and with a horizontal and vertical stabilizer. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight.
[0021] From U.S. Pat. No. 5,174,523 A (David E. H. Balmford) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a propeller with a flow guiding unit at the backside of the helicopter, and with two additional wings. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight and the torque compensation by means of the flow guiding unit.
[0022] From RU 2 089 456 A (Mikhail Il'ich Fefer) there is known an aircraft configuration which is designed as a helicopter with two wings which are arranged in the central region of the fuselage, wherein at the ends of said two wings there is rigidly arranged a main rotor, respectively. The wings are rigidly arranged at the aircraft fuselage below the respective main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two additional wings.
[0023] From U.S. Pat. No. 5,067,668 A (Daniel R. Zuck) there is known an aircraft designed as a helicopter with additional wings below the main rotor, wherein said wings are designed in a pivotable manner and thereby enable the torque compensation during the hover flight or low speed flight, and hence the tail rotor as a torque compensation is omitted. The propeller arranged at the tail is used exclusively for the generation of a thrust for the forward flight. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor.
[0024] From U.S. Pat. No. 4,928,907 (Daniel R. Zuck) there is known an aircraft designed as a helicopter with additional wings below the main rotor, wherein said wings are designed in a pivotable manner and thereby enable the torque compensation during the hover flight or low speed flight, and hence the tail rotor as a torque compensation is omitted. A propeller arranged at the tail is used exclusively for the generation of a thrust for the forward flight. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor.
[0025] From U.S. Pat. No. 4,691,877 A (Ralph M. Denning) or GB 2143483 (John Denman Sibley) there is known an aircraft which is designed as a helicopter with additional wings below the main rotor, and pivotable flaps are arranged at the wings, around which the exhaust gas of the afterburner from the main drive flows. The wings are rigidly connected to the aircraft fuselage. During take-off and landing as well as in the hovering state the vertical lift is generated by the main rotor and by the exhaust gas flow from the two afterburners which can also generate a torque compensation and an additional thrust in forward flight.
[0026] From U.S. Pat. No. 3,977,812 A (Wayne A. Hudgins) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, and with two additional wings. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight.
[0027] From CA 825 030 A (Nagatsu Teisuke) or U.S. Pat. No. 3,448,946 A (Nagatsu Teisuke) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, with a horizontal and vertical stabilizer, and optionally with two additional wings. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight.
[0028] From the publication by C. Silva and H. Yeo, Aeroflightdynamics Directorate, U.S. Army RDECOM, and W. Johnson, NASA Ames Research Center: “Design of a Slowed-Rotor Compound Helicopter for Future Joint Service Missions”, Aeromech Conference, San Francisco, Calif., January 2010, there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, with a horizontal and vertical stabilizer, and with two additional wings. The wings are rigidly arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight.
[0029] From U.S. Pat. No. 3,563,469 A (Daniel R. Zuck) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a propeller at the backside of the helicopter, with a horizontal and vertical stabilizer, and with two additional pivotable wings. The wings are arranged in a pivotable manner at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside propeller generates the thrust for the forward flight, the tail rotor generates the torque compensation.
[0030] From U.S. Pat. No. 3,241,791 A (F. N. Piasecki) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a ducted fan at the backside of the helicopter, with two additional wings arranged at the aircraft fuselage below the main rotor, and with a flow guiding unit at the output of the ducted fan. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The backside ducted fan with the flow guiding unit generates the thrust for the forward flight and the torque compensation.
[0031] From CA 700 587 A and U.S. Pat. No. 3,105,659 A (Richard G. Stutz) there is known an aircraft configuration which is designed as a helicopter with a main rotor, a tail rotor, a horizontal stabilizer, and with two additional rigid wings with aileron flaps and propellers. The wings are arranged at the aircraft fuselage below the main rotor. During take-off and landing as well as in the hovering state the vertical lift is generated exclusively by the main rotor, while in the forward flight an additional lift is generated by the two wings. The tail rotor generates the torque compensation, and the two propellers generate the thrust in forward flight.
[0032] All said known compound helicopter aircraft configurations which are designed with classical thrust generating means like propellers have the disadvantage that the vertical lift for the take-off and the landing as well as in the hovering state is exclusively or mainly generated by the main rotor, and that a correspondingly large main rotor diameter is required. In the forward flight the large main rotor produces the largest flow resistance and causes the largest energy loss. The additional drive units like propellers or ducted fans facilitate higher flying velocities and improved maneuverabilities, but with an increasing flying velocity the efficiency is reduced and the energy consumption is increased disproportionally.
[0033] The known compound helicopter aircraft configurations with cyclogyro rotors have the disadvantage that in the known lateral arrangement of the cyclogyro rotors the rotor discs and support elements influencing the aerodynamics are lacking and that the cyclic adjustment of the rotor blades has to be carried out by a rotating rotor shaft and can only be carried out from the side facing the aircraft fuselage, respectively, that with the known horizontal arrangement as a tail rotor no contribution to the generation of the thrust force in the flight direction can be made, and that the through-flow cross-section in the rotor is reduced massively by the helicopter structure, and that with the vertical arrangement as tail rotor no contribution to the vertical thrust generation can be made.
[0034] From U.S. Pat. No. 5,100,080 A, U.S. Pat. No. 5,265,827 A, and US 2007/0200029 A1 there are known aircrafts with cyclogyro rotors having adjustable rotor blades. Combinations with main rotors are not mentioned therein. Therefore, the advantages of a helicopter cannot be utilized.
SUMMARY OF THE INVENTION
[0035] The object of the present invention is to define a novel aircraft on the basis of a helicopter which avoids the above-described disadvantages without losing the additional benefits.
[0036] This object is achieved according to the invention in that the cyclogyro rotors are connected to the aircraft fuselage by means of a suspension device which holds the cyclogyro rotors at the outer border of the rotors, and that each cyclogyro rotor can be controlled individually and independently of the other, and that a torque compensation function of the main rotor can be carried out by the cyclogyro rotors.
[0037] In this case it is a helicopter which is equipped with two additional laterally arranged cyclogyro rotors which—independently of one another—can generate a thrust vector to be controllable in any direction in a plane substantially in parallel to the axis of rotation of the main rotor and to the longitudinal axis of the helicopter, and hence they can take over the torque compensation of the main rotor in all flying situations, they will supplement the vertical thrust of the main rotor in the vertical take-off, landing, and hovering state, they will support the secure transition from the hovering state into the forward flight, and they will generate the required thrust in forward flight. Due to the support of the vertical thrust of the main rotor, the main rotor can be made with a diameter smaller in size compared to the classical helicopter and all known compound (hybrid) helicopters so that in forward flight a better efficiency can be obtained or that with a comparable driving power a higher velocity can be achieved. A tail rotor in the classical sense is not necessary.
[0038] In this connection, two cyclogyro rotors are connected to the aircraft fuselage laterally by means of a support device or with supporting elements such that the thrust forces generated by the cyclogyro rotor can be introduced into the aircraft fuselage so that a substantially lighter construction can be achieved. Furthermore, by the cyclogyro rotors which can be controlled independently of the other the torque compensation is taken over so that no tail rotor is necessary which facilitates a further reduction in weight.
[0039] The connection of the suspension device to the aircraft fuselage can also be effected indirectly by means of other components.
[0040] The bilateral mounting of the cyclogyro rotors in the suspension device is of special importance, as it does not only facilitate a lighter and more robust construction, but also an adjustment of the rotor blades on both sides.
[0041] In a particular embodiment of the invention the offset adjustment devices required for the cyclic adjustment of the rotor blades are arranged on both sides of the cyclogyro rotor, whereby a light and robust construction is obtained which puts the least load on the critical rotor components. The introduction of the torque into the cyclogyro rotors takes place at the side of the cyclogyro rotor facing the aircraft fuselage.
[0042] By the preferred arrangement of the cyclogyro rotors below the main rotor, the main rotor can be reduced in size to a substantial degree, as the generation of the vertical thrust for the vertical take-off, landing, and for the hovering state is supported by two cyclogyro rotors arranged laterally at the aircraft fuselage below the main rotor. A cyclogyro rotor generates a thrust vector that can be controlled in a plane perpendicular to the axis of rotation of the rotor in any direction and can be adjusted continuously from 0 up to a maximum value by changing a cyclic pitch angle of the rotating rotor blades as a function of the displacement of an offset position within the rotating cyclogyro rotor. By the lateral arrangement of such rotors at one side of the aircraft fuselage, respectively, and by the unlimited change of direction of the thrust vectors, said rotors furthermore generate the torque compensation of the main rotor so that in this configuration no tail rotor is required. The configuration according to the invention facilitates a vertical take-off helicopter which shows a lower energy consumption at the same carrying capacity, which has a smaller main rotor diameter and thus can take off and land also on a smaller space, which does not require any tail rotors, and which achieves a higher flying velocity with a comparatively lower energy consumption. The compound helicopter according to the invention also has the potential of a larger range with the same fuel load. A further advantage is the higher agility in almost all flight phases.
[0043] It is preferred if the suspension device is designed as wings or airfoils in order to generate a lift in the forward flight. Thereby, on the one hand, the load on the main rotor can be reduced and, on the other hand, the maximum velocity can be increased, as the main rotor can be operated with a reduced speed.
[0044] In this connection it is particularly favorable if the suspension device is arranged above the cyclogyro rotors. In this way, in the forward flight an improved flow against the cyclogyro rotors can be achieved. In order to improve the effect of the main rotor onto the cyclogyro rotors it can in particular be provided that the suspension device has a recess directly above the cyclogyro rotors.
[0045] Preferably there is provided a horizontal and vertical stabilizer for the stabilization. This means in particular that no separate airscrew is provided in order to manage the torque compensation, which is also not required due to the design according to the invention.
[0046] A particular embodiment of the invention provides that the cyclogyro rotors are connected to the drive of the main rotor by means of a gear. This means that the speeds of the main rotor and of the cyclogyro rotors always will be in a constant proportion to each other. The respectively required thrust will be adjusted by the adjustment of the rotor blades. This enables a very simple drive.
[0047] Alternatively it can be provided that the cyclogyro rotors have a drive which is independent of the main rotor, wherein said drive can be electrical, hydraulic, or can be implemented as an individual drive unit. Thereby the thrust can be varied within particularly broad limits.
[0048] In a particularly advantageous embodiment the aircraft does not have any tail rotor. Thereby the weight can be reduced and the constructional expenditure is reduced.
BRIEF DESCRIPTION OF THE FIGURES
[0049] The invention will now be described in detail by means of FIGS. 1 to 8, wherein:
[0050] FIG. 1 shows a compound helicopter according to the invention in a diagonal view from above;
[0051] FIG. 2 shows the helicopter of FIG. 1 in a view from the front;
[0052] FIG. 3 shows the helicopter of FIG. 1 in a lateral view;
[0053] FIG. 4 shows the helicopter of FIG. 1 in a plan view;
[0054] FIG. 5 shows a cyclogyro rotor in a diagonal view;
[0055] FIG. 6 shows the cyclogyro rotor in a lateral view;
[0056] FIG. 7 shows the cyclogyro rotor in a view from the front; and
[0057] FIG. 8 shows an offset adjustment device in detail.
DETAILED DESCRIPTION OF THE INVENTION
[0058] FIG. 1 shows an aircraft according to the invention, namely a compound helicopter, in a diagonal view from above, consisting of the aircraft fuselage 1, the main rotor 2, the laterally arranged cyclogyro rotors 3 and 3′, the suspension 4 and 4′ of the cyclogyro rotors, the outer mounting or bearing 5′, and the outer offset adjustment device 11′, and the horizontal and vertical stabilizer 6, 6′, 7, 7′, and the recess 20 in the suspension 4 and 4′.
[0059] FIG. 2 shows the compound helicopter according to the invention in a front view, with the two laterally arranged cyclogyro rotors 3 and 3′, their suspension 4 and 4′, wherein the suspension 4 and 4′ is also designed as a wing or airfoil or as a component having the function of a wing or airfoil. In the central region there is provided a recess 20 which facilitates a passage of air downwards. The suspension 4, 4′ is fixed on the one hand at the aircraft fuselage 1 and on the other hand at the outside of the cyclogyro rotors 3, 3′ and holds them.
[0060] An offset adjustment device 11 and 11′ for the adjustment of the rotor blades 9 is arranged at the outside of the cyclogyro rotors 3, 3′, wherein the two offset adjustment devices facing the aircraft fuselage 1 of the helicopter are not visible. Thereby it is possible to perform the cyclic adjustment of the rotor blades 9 from two sides and to provide the drive of the rotor 3, 3′ from the side facing the aircraft fuselage 1 of the helicopter. It is provided that the cyclogyro rotors 3, 3′ have a length in the axial direction (e.g., a distance from the aircraft fuselage 1 to the outer border) which corresponds approximately to the diameter of the cyclogyro rotors 3, 3′ and preferably lies between 80% and 120% of the diameter.
[0061] FIG. 3 shows the compound helicopter according to the invention in a lateral view, with the laterally arranged cyclogyro rotor 3′, its suspension 4′, wherein the suspension can also be implemented as a wing or as a component having the function of a wing, the outer rotor mounting or bearing 5′, and the outer offset adjustment device 11′, and the vertical stabilizer 6.
[0062] From FIG. 4 there becomes evident in particular the horizontal and vertical stabilizer 6, 6′, 7, 7′.
[0063] FIG. 5 shows the right-hand side cyclogyro rotor 3 of FIG. 2 in a diagonal view, consisting substantially of the rotor shaft 10, the rotor blades 9 (preferably three to six), the two rotor disks 8 and 8′ with integrated rotor blade bearing or mounting, the lateral offset adjustment device 11 facing away from the helicopter aircraft fuselage, for influencing the cyclic pitch angle of the rotor blades and the direction of the thrust vector 12 which can be controlled in a plane 15 perpendicular to the axis of rotation 10 of the rotor into any direction and any size, if the cyclogyro rotor 3 is kept in rotation with a corresponding speed according to the rotational direction 14.
[0064] FIG. 6 shows the cyclogyro rotor 3 in a lateral view, wherein by the angle φ the direction 13 of the thrust vector 12 is indicated and by Ω the direction of rotation 14 of the cyclogyro rotor is indicated.
[0065] FIG. 7 shows the right-hand side cyclogyro rotor 3 of FIG. 2 in a lateral view, consisting substantially of the two rotor disks 8 and 8′, the rotor shaft 10, the rotor blades 9 (preferably 3 to 6), the lateral offset adjustment device 11 facing away from the aircraft fuselage 1 of the helicopter, and the offset unit 19 facing the aircraft fuselage 1 of the helicopter, for influencing the cyclic pitch angle of the rotor blades and the direction of the thrust vector.
[0066] FIG. 8 shows the cyclic rotor blade setting devices 16 which are connected in the rotor disks 8 to the offset adjustment device 11. By displacing the central offset point 17 within a circular area 18, in accordance with the distance and the direction of the offset point 17 from the axis of rotation 10 of the rotor the size of the thrust vector and the direction of the thrust vector will be defined.
WO2022243559 -- Aircraft [ PDF ]
Inventor(s): HOFREITHER KLEMENS [AT]; KINAST LUKAS [AT] +
Applicant(s): CYCLOTECH GMBH [AT] +
The invention relates to an aircraft (100) comprising an aircraft body (120) which defines a longitudinal direction, a vertical direction and a transverse direction, and at least two drive devices (1) which can rotate about a respective associated rotational axis (5) in order to generate a respective associated thrust vector, wherein a first number of the drive devices are arranged along a first straight line running parallel to the transverse direction, and a second number of drive devices are arranged along a second straight line running parallel to the transverse direction, the first straight line is spaced apart from the second straight line, and the centre of gravity of the aircraft is positioned between the first straight line and the second straight line. The aircraft is designed to perform a hovering flight, such that, in the hovering flight, each of the associated rotational axes (5) is orientated substantially in the transverse direction of the aircraft body, and each of the at least two drive devices rotates about the respective associated rotational axis substantially in the same rotational direction. The invention also relates to aircraft configurations with further orientations of the rotational axes.
The invention relates to an aircraft and methods for producing and controlling the aircraft. In particular, the invention relates to an aircraft that can achieve a stable hovering flight with drive devices rotating in the same direction, in particular cyclogyro rotors. Aircraft that use cyclogyro rotors as propulsion devices are called cyclogyros. Cyclogyros, like helicopters, are also so-called Vertical take-off and landing vehicles (VTOL vehicles) are aircraft that are capable of taking off and landing vertically without a runway. A cyclogyro rotor is based on the principle of generating thrust with rotating wings, which are then called rotor blades. In contrast to classic rotating blades, such as those used in the propulsion system of a helicopter, the rotation axis of the blades of a cyclogyro rotor is aligned parallel to the longitudinal axis of the wings / rotor blades.
The thrust direction of the entire cyclogyro rotor is normal to the axis of rotation. In stationary operation, such as hovering or forward flight at constant speed, all rotor blades of the cyclogyro rotor should ideally be aligned as best as possible to the flow direction at all times in order to make a maximum contribution to the total thrust with the minimum required propulsion power. The maximum inclination of the rotor blades relative to the flow direction directly influences the amount of thrust generated. Due to the rotation of the rotor, the inclination of each rotor blade must be continuously changed during one revolution. Each rotor blade of a cyclogyro rotor thus undergoes a periodic change of the angle of inclination. This periodic change in the angle of inclination is called pitch movement. Different pitch mechanisms are known to generate the pitch movement. For example, each rotor blade can be connected to an eccentric bearing axis via one or more connecting rods.
The resulting pitch movement of a rotor blade repeats itself cyclically with each rotor revolution. Various designs of drive devices for cyclogyros are described in the European patent applications published under numbers EP 3548378 A1 and EP 3715249 A1. The periodic adjustment of the rotor blades creates a thrust vector normal to the rotor's axis of rotation. With the help of an offset device, the periodic rotor blade adjustment is changed, and thus the thrust vector can be rotated in the entire plane which is normal to the axis of rotation of the rotor (thrust vector control). In addition to the thrust vector, the rotor generates a torque around the axis of rotation opposite to the direction of rotation of the rotor resulting from the tangential components of the air forces acting on the rotor blades, namely the lift and drag forces. If air flows onto the rotor from the outside, the aerodynamic properties and thus the properties of the generated thrust vector change.
When the rotor is in forward flight, it is actively supplied with air from the front. The changed properties can be approximately explained by the Magnus effect. This states: “A rotating round body in a flow experiences a transverse force normal to the flow direction. “ The direction of the transverse force depends on the direction of rotation of the body or in this case: the cyclogyro rotor. For example, from the article by I.S. Hwang et al.: “Development of a Four-Rotor Cyclocopter” from Journal of Aircraft, Vol. 45, No. 6, November–December 2008, pages 2151 ff. and the article by M. Benedict et al.: “Experimental Optimization of MAV-Scale Cycloidal Rotor Performance” from Journal of the American Helicopter Society 56, 022005 (2011) known aircraft or However, cyclogyros rotate rotors in opposite directions while maintaining constant airflow. In this case, i.e. when the rotors rotate in opposite directions, the transverse forces of the rotors caused by the Magnus effect do not act in the same direction and can thus reduce the overall thrust or increase the power requirement while the same lift force is required.
At higher forward speeds and in the opposite direction of rotation, it is therefore possible that the negative impact of the Magnus effect can no longer be compensated by the rotor. As a result, the aircraft is no longer capable of flying and the rotor cannot be used as a lift-generating component. The object of the present invention is therefore to provide an aircraft which is capable of maintaining a stable flight attitude even at high speeds in forward flight. This object is achieved by the aircraft having the features according to claim 1, by the aircraft having the features according to claim 5, the methods for producing an aircraft according to claims 17 and 18, respectively, and the methods for controlling an aircraft according to claims 19 and 20, respectively. Advantageous embodiments of the present invention are specified in subclaims 2 to 4, 6 to 16 and 21 to 24. According to a first aspect of the invention, an aircraft is provided comprising the following components: an aircraft body defining a longitudinal direction, a vertical direction and a transverse direction, wherein the longitudinal direction corresponds to the direction from the tail to the nose of the aircraft, the vertical direction corresponds to the direction of gravity when the aircraft is resting on the ground, and the transverse direction is perpendicular to the longitudinal direction and the vertical direction, and at least two drive devices which are rotatable about a respective associated axis of rotation in order to generate a respective associated thrust vector.
A first number of drive devices is arranged along a first straight line running parallel to the transverse direction, and a second number of drive devices is arranged along a second straight line running parallel to the transverse direction. The first line is spaced from the second line, and the center of mass of the aircraft is positioned longitudinally between the first line and the second line. The aircraft is designed to perform a hover flight in which all forces acting on the aircraft acting forces and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear, due to the fact that, in hovering flight, each of the associated axes of rotation is aligned essentially in the transverse direction of the aircraft body, and each of the at least two drive devices rotates essentially in the same direction of rotation about the respectively associated axis of rotation. According to the invention, an axis of rotation is aligned substantially in the transverse direction of the aircraft body if the angle enclosed between the axis of rotation and an axis running in the transverse direction and intersecting the axis of rotation is less than 45°, preferably less than 30°, particularly preferably less than 15°.
In the sense of the invention, it is therefore not necessary that all axes of rotation are mathematically exactly parallel during hovering. It may even be expedient if the angle between an axis of rotation and an axis which runs in the transverse direction and intersects the axis of rotation is in the range between 5° and 30°, particularly preferably between 10° and 20°. Furthermore, according to the invention, the drive devices rotate essentially in the same direction of rotation if the scalar product of the vector of the angular velocity of a specific drive device and a fixed but arbitrary vector pointing in the transverse direction has the same sign for all drive devices. This means that in order to check that all drive devices under consideration or each of the drive devices under consideration rotate essentially in the same direction of rotation, a vector in the transverse direction is first fixed. Subsequently, the scalar product of the angular velocity vector of a first drive device and the fixed vector is calculated; then, for a second drive device, the scalar product of the angular velocity vector of a second drive device and the fixed vector; etc.
Finally, only the signs (plus or minus) of the scalar products calculated in this way are compared. If all signs are the same, the drive devices under consideration rotate or each of the drive devices under consideration rotates essentially in the same direction of rotation within the meaning of the invention. In the sense of the invention, it is therefore neither necessary that all axes of rotation are mathematically exactly parallel during hovering, nor that all Drive devices rotate around the axis of rotation with the same rotational or (magnitude) angular velocity. The fact that the aircraft is designed to perform a hover flight with the propulsion devices rotating essentially in the same direction results in a reduction in the power consumption of the propulsion devices. Put simply, the Magnus effect occurring according to the invention replaces part of the thrust of the propulsion devices and thus reduces the power requirement in forward flight compared to hovering flight.
Because more residual power remains for the propulsion systems in forward flight, the agility of the aircraft in particular increases in forward flight. The Magnus effect states that a rotating round body in a flow experiences a transverse force normal to the flow direction. In the case of the drive devices according to the invention, which rotate essentially in the same direction, this effect can generate an additional thrust vector or an additional thrust force in the vertical direction. This increases the overall lift force of the propulsion devices. The Magnus effect replaces part of the thrust required by the propulsion system and thus reduces the power required in forward flight compared to hovering. If the rotor is now in forward flight, it is actively supplied with air from the front. In the configuration according to the invention, in which the drive devices rotate substantially in the same direction, the additional transverse force of the Magnus effect acts substantially in the same direction as the thrust of the drive devices when the flow remains constant and thus increases the total thrust or reduces the power requirement while the same lift force is required.
In forward flight, especially at higher forward speeds and essentially the same direction of rotation, it is therefore possible that the positive impact of the Magnus effect requires a lower power and/or rotation speed of the propulsion devices in order to keep the aircraft in a stable flight attitude. Particularly preferably, the aircraft is further designed such that, during hovering, the center of mass of the aircraft is positioned in such a way that all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft are in the Essentially disappear when one or more of the propulsion devices generate a specific predetermined thrust vector assigned to them. This instruction is subject to the restriction that the longitudinal centre of mass of the aircraft must be within a range determined by the ability of the aircraft to hover when one or more of the propulsion devices are driven at maximum thrust or maximum thrust vector.
In other words, if the center of mass is within the said area, the propulsion devices are able to generate corresponding thrust vectors so that the aircraft can perform hovering. In hovering flight, the airflow velocity is generally lower than in forward flight. By specifying the thrust vectors of the propulsion devices for hovering flight for the aircraft according to the invention and determining the position of the center of mass for hovering flight, it is ensured that a stable flight attitude is also possible in forward flight. As stated above, the positive effect according to the invention, which is caused by the Magnus effect, is greater the higher the flow velocity is. Therefore, the inventive configuration of the aircraft in hover flight ensures that the aircraft can assume a stable flight attitude, especially in forward flight, because in forward flight the Magnus effect leads to a greater increase in the thrust vector than in the case of hover flight.
When designing and configuring an aircraft according to the invention with propulsion devices, all forces and torques of the propulsion devices must be taken into account. Basically, the thrust force or thrust vector is used to generate the required lift force and/or to control the flight attitude of the aircraft. For this purpose, the aircraft expediently comprises a thrust vector control which regulates the required thrust force or thrust vectors in hovering flight and/or in forward flight. Each of the drive devices according to the invention generates a torque opposite to the direction of rotation. This torque around the axis of rotation opposite to the direction of rotation of the drive device results from tangential air forces caused, among other things, by air resistance. In order to maintain a constant rotational speed, the drive device must therefore generate a (drive) torque that corresponds to the counteracts the torque resulting from tangential air forces.
However, in order for the propulsion device to be able to generate such a (drive) torque during the flight phase, an additional torque is required, which the aircraft body must generate (according to the principle of actio = reactio) in order to “support” the propulsion device in the air. In order to maintain a constant rotational speed against the air forces, this latter torque is (neglecting dissipative effects) approximately equal in magnitude to the torque generated by the tangential air forces, and also points in the same direction as the latter. Since the torque generated by the air forces counteracts the direction of rotation of the propulsion device, the torque applied by the aircraft body also counteracts the direction of rotation of the propulsion device. Assuming that the torque due to the air forces and that of the propulsion device are essentially equal in magnitude but oppositely directed, the net torque remaining is the torque applied by the aircraft body due to the rotation of the propulsion device.
According to the invention, this torque or these torques are compensated by positioning the center of mass of the aircraft in such a way that, taking into account the thrust vectors assigned to the respective drive devices and predetermined, all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear during hovering flight. Because, according to the invention, the drive devices rotate essentially in the same direction of rotation, the torques of all of these drive devices caused by the aircraft body as described above also act essentially in the same direction. The torques therefore add up and do not cancel each other out. In order to achieve a stable flight position in hovering and forward flight, the balance of all forces and torques acting on the aircraft must be achieved. The calculation is carried out using the momentum and angular momentum theorem.
The momentum theorem is: <img file="imgf000010_0001.tif" frnum="0001" he="11" id="imgf000010_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0001" wi="26"/> where m is the mass of the aircraft, <img file="imgf000010_0003.tif" frnum="0001" he="7" id="imgf000010_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0002" wi="7"/> is the acceleration vector of the center of mass of the aircraft and F is the force vector acting on the aircraft. The angular momentum theorem states <img file="imgf000010_0002.tif" frnum="0001" he="9" id="imgf000010_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0003" wi="23"/> where M<sub>s</sub> is the temporal change of the angular momentum vector (angular momentum vector) and M<img file="imgf000010_0002.tif" frnum="0001" he="9" id="imgf000010_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0003" wi="23"/> is the torque vector acting on the aircraft.
If a stable flight attitude is required (hovering, constant speed in forward flight, etc.), the acceleration vector <img file="imgf000010_0005.tif" frnum="0001" he="7" id="imgf000010_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0004" wi="5"/> and the temporal change of the angular momentum vector <img file="imgf000010_0004.tif" frnum="0001" he="7" id="imgf000010_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0005" wi="6"/> must be zero. Thus, both the sum of all external forces (F) and the sum of all torques around the center of mass (M<sub>s</sub>) must be zero. The forces acting on the aircraft in hovering flight are gravity and the thrust of the propulsion systems.
The torques acting with respect to the centre of mass of the aircraft are the torques generated by the thrust vectors of the propulsion devices mounted at appropriate distances from the centre of mass of the aircraft, as well as the (support) torques generated by the aircraft body, all of which point substantially in the same direction. The force and torque equilibrium can thus be achieved by balancing the thrust forces or Thrust vectors of the propulsion devices and their distances from the center of mass of the aircraft are selected accordingly. Preferably, the first number of drive devices is arranged in a front region of the aircraft with respect to the longitudinal direction, and the second number of drive devices is arranged in a rear region of the aircraft with respect to the longitudinal direction. Preferably, the aircraft comprises three propulsion devices. Particularly preferably, the aircraft comprises four drive devices, wherein two of the drive devices are arranged in a front region of the aircraft with respect to the longitudinal direction, and two further drive devices with respect to the
longitudinally in a rear area of the aircraft. The total length of the aircraft is measured longitudinally. To simplify the description of areas of the aircraft, the frontmost part of the aircraft is assigned the relative longitudinal coordinate 0 and the rearmost part of the aircraft is assigned the relative longitudinal coordinate 100%. In this convention, the front area is defined as corresponding to the (longitudinal) range from 0 to 40% of the total length of the aircraft, and the rear area is defined as corresponding to the (longitudinal) range from 60% to 100% of the total length of the aircraft. Furthermore, it is expedient if the two drive devices arranged in the front area are located on a common straight line which is aligned parallel to the transverse direction. It is also expedient if the two drive devices arranged in the rear area are located on a common straight line that is aligned parallel to the transverse direction. Advantageously, the drive devices in the front region are arranged along the first straight line which runs parallel to the transverse direction, and the drive devices in the rear region are arranged along the second straight line which runs parallel to the transverse direction.
The centre of mass of the aircraft, when it is hovering, is positioned longitudinally at a distance l<sub>1</sub>from the straight line along which the propulsion devices are arranged in the front area, where <img file="imgf000011_0001.tif" frnum="0001" he="30" id="imgf000011_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0006" wi="136"/> R<sub>min</sub>is a minimum permissible ratio between the thrust vectors of the propulsion devices arranged in the front area, on the one hand, and the thrust vectors of the propulsion devices arranged in the rear area, on the other hand, R<sub>max</sub>is a maximum permissible ratio between the thrust vectors of the propulsion devices arranged in the front area, on the one hand, and the thrust vectors of the propulsion devices arranged in the rear area, on the other hand, l is the distance between the first straight line and the second straight line, a<sub>1</sub>is an index for the propulsion devices arranged in the front area, and a<sub>2</sub>is an index for the propulsion devices arranged in the rear area.
For practical purposes, the aircraft is further designed so that the associated axes of rotation are aligned parallel during hovering flight. Finally, it should be pointed out that the invention does not exclude the possibility that the aircraft, in addition to the at least two drive devices contributing to the effect according to the invention, comprises further drive devices which do not rotate essentially in the same direction of rotation. According to a second aspect of the invention, an aircraft is provided which comprises an aircraft fuselage and at least three drive devices which are mounted around the aircraft fuselage and which are rotatable about a respective associated axis of rotation in order to generate a respective associated thrust vector. The aircraft is designed to perform a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear, in that in hover flight the associated axes of rotation of two of the at least three drive devices are aligned essentially in a first direction, and the associated axis of rotation of a further one of the at least three drive devices is aligned essentially in a second direction, wherein the first direction is not parallel to the second direction, and each of the two drive devices with axes of rotation aligned in the first direction in hover flight rotates essentially in the same direction of rotation about the respectively associated axis of rotation.
For the inventive understanding of the terms “substantially aligned in a first/second direction” and “substantially rotating in the same direction”, reference is made to the first aspect of the invention; the definitions given there apply accordingly to the second aspect. The first direction is not parallel to the second direction if a (reference) axis pointing in the first direction is not parallel to a (reference) axis pointing in the second direction. Preferably, the angle between the first and second direction is in the range of 30° to 110°, preferably in the range of 40° to 100°, particularly preferably in the range of 60° to 95°. Preferably, the at least three drive devices are mounted around the aircraft fuselage substantially in one plane. It is practical if the aircraft fuselage lies in the plane, i.e. the plane intersects the aircraft fuselage. Furthermore, it is advantageous if the first direction and the second direction are in the plane.
Here, “substantially supported in one plane” means that the drive devices or their support points do not have to be contained in one plane in exactly the same way. It is also according to the invention if one or more of the drive devices are pivoted out of the plane and/or the drive devices are vertically offset with respect to the plane. Advantageously, the vertical offset is limited by the vertical extension of the aircraft fuselage, i.e., the drive devices are advantageously mounted in such a way that the axes of rotation of the drive devices are contained in the spatial region formed between two horizontal planes which touch the aircraft fuselage and are spaced apart from one another by the vertical extension of the aircraft fuselage. The vertical extension is related to the direction of gravity when the aircraft is resting on the (flat) ground. Preferably, each of the axes of rotation of the two of the at least three drive devices aligned substantially in the first direction is aligned such that it is substantially parallel to a straight line running through the two drive devices.
It is useful if the straight line is placed through the geometric centers (the term is explained below) or bearing points of the drive devices. According to the invention, an axis of rotation is substantially parallel to a straight line if the angle enclosed between the axis of rotation and the straight line is less than 45°, preferably less than 30°, particularly preferably less than 15°. Particularly preferably, the aircraft according to the second aspect of the invention comprises at least four drive devices which are mounted around the aircraft fuselage and which are rotatable about a respective associated axis of rotation in order to generate a respective associated thrust vector. The aircraft is designed to perform hovering flight in that, during hovering flight, the associated axes of rotation of two of the at least four drive devices are directed essentially in the first direction are aligned, and the associated axes of rotation of two further ones of the at least four drive devices are aligned substantially in the second direction, wherein each of the two drive devices with axes of rotation aligned in the first direction during hovering rotates substantially in the same direction of rotation about the respectively associated axis of rotation, and/or each of the two drive devices with axes of rotation aligned in the second direction during hovering rotates substantially in the same direction about the respectively associated axis of rotation.
The advantages that the aircraft according to the second aspect of the invention brings with it over the prior art basically correspond to those that have already been described in connection with the aircraft of the first aspect of the invention; in order to avoid repetition, reference is therefore first made to the explanations therein, in particular to the utilization of the positive contribution of the Magnus effect in the case of drive devices rotating in the same direction. In connection with the latter contributions of the Magnus effect, when arranging the propulsion systems around the aircraft fuselage – also referred to below as a “star-shaped” arrangement – it must be taken into account that in forward flight, as a rule, only a part of the propulsion systems is exposed to air flow in the direction of flight. Thus, the Magnus effect in forward flight is most pronounced in those propulsion devices whose axes of rotation are aligned essentially perpendicular to the direction of flight due to the essentially equal rotational rotation.
That is, in the arrangement of the drive devices according to the second aspect of the invention, it is sufficient if the aircraft is configured such that, in hovering flight, each of the two drive devices with axes of rotation aligned in the first direction in hovering flight rotates substantially in the same direction of rotation about the respectively assigned axis of rotation, or, in the case of at least four drive devices, each of the two drive devices with axes of rotation aligned in the second direction in hovering flight rotates substantially in the same direction of rotation about the respectively assigned axis of rotation. In this case, it is possible that the two drive devices, which do not rotate in substantially the same direction, rotate in opposite directions. If these two drive devices rotate in opposite directions, the torque cancels itself out. However, it is particularly advantageous that the aircraft, if it comprises at least four drive devices, is configured such that in hovering flight each of the two drive devices with axes of rotation aligned in the first direction
rotates substantially in the same direction of rotation about the respectively associated axis of rotation, and each of the two drive devices with axes of rotation aligned in the second direction during hovering rotates substantially in the same direction of rotation about the respectively associated axis of rotation. This ensures that the aircraft can exploit the positive effect of the Magnus effect during forward flight in both the first and second directions. This makes the aircraft more flexible and stable when changing direction. Particularly preferably, the aircraft is further designed such that, in hovering flight, the center of mass of the aircraft is positioned in such a way that all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear when one or more of the drive devices generate a specific predetermined thrust vector assigned to them.
This instruction imposes the restriction that the centre of mass of the aircraft must be within a range determined by the ability of the aircraft to hover when one or more of the propulsion devices are driven at maximum thrust or maximum thrust vector. In other words, if the center of mass is within the said area, the propulsion devices are able to generate corresponding thrust vectors so that the aircraft can perform hovering. It is preferred if each of the axes of rotation of the two of the at least four drive devices aligned substantially in the first direction is aligned such that it runs substantially parallel to a straight line running through the two drive devices. It is also preferred if each of the axes of rotation of the two further drive devices of the at least four drive devices, which axes are aligned substantially in the second direction, is aligned such that it runs substantially parallel to a straight line which runs through these two further drive devices.
It is useful if the straight lines are placed through the geometric centers or bearing points of the drive devices. As in the first aspect of the invention, the compensation of the force(s) generated by the drive devices rotating essentially in the same direction is carried out torque or Torques are reduced according to the invention in that the center of mass of the aircraft is positioned in such a way that, taking into account the predetermined thrust vectors assigned to the drive devices in each case, all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear. In order to achieve a stable flight position in hovering and forward flight, the balance of all forces and torques acting on the aircraft must be achieved. The calculation is carried out using the momentum and angular momentum theorems, which have already been stated and described in connection with the first aspect of the invention.
The statements made there apply here accordingly, and this will be explained in more detail below. It is advantageous if three propulsion devices are arranged around the aircraft fuselage in such a way that they form the corners of a triangle, preferably an equilateral triangle. Conveniently, the aircraft fuselage is located in the geometric center of the triangle. The first direction is defined by a straight line on which two of the three drive devices lie; the second direction is essentially perpendicular to the first direction. Furthermore, the axis of rotation of each of the two drive devices lying on the straight line pointing in the first direction encloses an angle with said straight line which lies in the range between 0° and 45°, expediently between 0° and 30°. The geometric center corresponds to the average of all points within the triangle (i.e. the average over the area of the triangle with constant density). If the angle between the axis of rotation(s) and the straight line pointing in the first direction is set to 30°, the axis of rotation(s) of the drive devices point(s) towards (or away from) the geometric centre.
However, the angle can also be selected differently for each of the drive devices. It is useful if the straight line is placed through the geometric centers or bearing points of the drive devices. It is advantageous if n propulsion devices are arranged around the aircraft fuselage in such a way that they form the corners of an n-gon, n > 3, expediently the corners of a regular n-gon, n > 3. Conveniently, the aircraft fuselage is located in the geometric center of the n-gon. The first direction is defined by a first straight line on which two of the n drive devices lie; the second direction is defined by a second straight line is defined on which two more of the n drive devices lie. The axis of rotation of each of the two drive devices lying on the first straight line pointing in the first direction encloses an angle with the first straight line which lies in the range between 0° and 45°, expediently between 0° and 30°, expediently in the range between 0° and 20°, particularly preferably in the range between 0° and 18°.
The rotation axes of different drive devices can enclose different angles with the first straight line. It is also expedient if the axis of rotation of each of the two drive devices lying on the second straight line pointing in the second direction encloses an angle with the second straight line which lies in the range between 0° and 45°, expediently between 0° and 30°, expediently in the range between 0° and 20°, particularly preferably in the range between 0° and 18°. The rotation axes of different drive devices can enclose different angles with the second straight line. If the angles are chosen as mentioned above, it is possible that the axes of rotation of the drive devices point towards the geometric center of the n-gon (or away from it). This particularly preferably means that the aircraft comprises 3, 4, 5, 6, 7, 8, ... drive devices which are arranged around the aircraft fuselage in such a way that they form the corners of an equilateral triangle, a square, a regular 5-, 6-, 7-gon, or regular 8-gon, etc.
The aircraft fuselage is expediently positioned essentially in the center of the n-gon, whereby the geometric center, but not the center of mass, of the n-gon is taken into account; because according to the invention the center of mass of the aircraft does not necessarily have to coincide with the geometric center (geometric center of gravity). The geometric center of an n-gon is defined according to the geometric center of the triangle. It is useful to have n = 2j, j > 1. Then it is further expedient that the aircraft fuselage is located between two opposite propulsion devices of the regular 2j-gon. In this case, it is advantageous if the two axes of rotation assigned to specific opposite drive devices each extend essentially in the direction defined by a straight line on which the two specific opposing drive devices lie. Furthermore, it is advantageous if the aircraft is designed to perform hovering flight in that, during hovering flight, two opposing drive devices rotate essentially in the same direction about their associated axis of rotation.
In this case, j directions according to the invention can then be defined. Advantageously, the angle between the first straight line and the second straight line is in the range between 60° and 100°, preferably between 60° and 90°, particularly preferably between 70° and 90°, particularly preferably between 72° and 90°. As shown later, for a regular (2j + 1)-gon, j > 1, it is particularly advantageous to choose the first line and the second line (or corresponding directions) such that the angle between the first line and the second line is 90° (1 – 1/(2j + 1)). For an (arbitrary) (2j + 1)-gon, a particularly preferred range for the angle between the first and second lines is given by: [90°Â (1 – 1/(2j + 1)); 90°]. If the angles between the axes of rotation of the drive devices arranged along the first straight line and the first straight line are in the range [0°; 90°/(2j + 1)], and/or the angles between the axes of rotation of the drive devices arranged along the second straight line and the second straight line are in the range [0°; 90°/(2j + 1)], configurations can be implemented in which the axes of rotation of the drive devices point in the direction of the geometric center of the (2j + 1) corner (or away from it).
In the case of a regular 2j-gon, j > 1, it is convenient to choose the first line and the second line so that they enclose an angle of 90° – 90°/(2j) Â (2j mod 4). Then the first and second lines each pass through the geometric center of the 2j-gon. For an (arbitrary) 2j-gon, a particularly preferred range for the angle between the first and second lines is given by: [90° – 90°/j; 90°]. If the first straight line and the second straight line are determined such that the angle between them is in the range [60°; 90°], and the angles between the axes of rotation of the drive devices arranged along the first straight line and the first straight line are in the range [0°; 30°], and/or the angles between the axes of rotation of the drive devices arranged along the second straight line and the second straight line are in the range [0°; 30°], the propulsion devices are arranged in an (arbitrary) regular n-gon (n > 2) around the aircraft fuselage, so that the axes of rotation of the propulsion devices are aligned towards the geometric center (or away from it).
If n > 3 is to be considered, it is sufficient if the angle between the axis of rotation of a drive device and the first or second straight line passing through it lies in the range [0°; 18°]. Conveniently, the second direction is substantially perpendicular, particularly preferably: perpendicular, to the first direction, and two of the at least four drive devices are arranged along the first direction, and the two further of the at least four drive devices are arranged along the second direction which is substantially perpendicular to the first direction. This is an example where the propulsion devices can be arranged around the aircraft fuselage at the corners of a square. Preferably, the center of mass of the aircraft, when it is hovering, is positioned in the first direction at a distance l<sub>34</sub>from a straight line along which the drive devices are arranged in the second direction, with <img file="imgf000019_0001.tif" frnum="0001" he="37" id="imgf000019_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0007" wi="92"/> where R<sub>min</sub>is a minimum permissible ratio between the thrust vectors of the drive devices arranged along the first direction, R<sub>max</sub>is a maximum permissible ratio between the thrust vectors of the drive devices arranged along the first direction, a<sub>34</sub>is an index for the drive devices arranged along the second direction, and l is the distance between the geometric centers of the drive devices arranged in the first direction.
Preferably, the center of mass of the aircraft, when it is hovering, is positioned in the second direction at a distance l<sub>12</sub> from a straight line along which the drive devices are arranged in the first direction, with <img file="imgf000020_0001.tif" frnum="0001" he="38" id="imgf000020_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0008" wi="97"/> wherein <img file="imgf000020_0002.tif" frnum="0001" he="8" id="imgf000020_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0009" wi="15"/> is a minimum permissible ratio between the thrust vectors of the drive devices arranged along the second direction, a<img file="imgf000020_0003.tif" frnum="0001" he="8" id="imgf000020_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0010" wi="15"/> is a maximum permissible ratio between the thrust vectors of the drive devices arranged along the second direction, a<sub>12</sub> is an index for the drive devices arranged along the first direction, and l<sup>'</sup> is the distance between the geometric centers of the drive devices arranged in the second direction.
It may be advantageous for both the aircraft according to the first aspect and that according to the second aspect to carry out hovering flight with approximately equal assigned specific predetermined thrust vectors. Likewise, it may be advantageous in each of the aircraft of the first or second aspect if it further comprises a displacement device with which the center of mass of the aircraft can be displaced. For this purpose, the aircraft expediently further comprises a fuel tank for supplying the drive devices with fuel and/or a battery for supplying the drive devices with electrical energy, wherein the displacement device is designed to displace fuel from the fuel tank or the battery within the aircraft in order to position the center of mass in such a way that the aircraft performs hovering flight when one or more of the drive devices generate the respectively associated specific predetermined thrust vector.
The center of mass of the aircraft can therefore be shifted dynamically. The advantage is that the center of gravity of the aircraft can be adjusted to various flight positions can be optimally adapted accordingly. The shift in the center of mass can be accomplished by aircraft control. Preferably, the aircraft according to the first or second aspect comprise a thrust vector control to individually control the thrust vectors of the propulsion devices. According to a third aspect of the invention, a method for producing an aircraft according to the first aspect of the invention is provided, comprising the following steps: - positioning the center of mass of the aircraft in such a way that one or more of the drive devices generate a specific predetermined thrust vector assigned to them in each case, so that the aircraft performs a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear when - each of the associated axes of rotation is aligned essentially in the transverse direction of the aircraft body, and - each of the at least two drive devices rotates essentially in the same direction of rotation about the respectively associated axis of rotation.
According to a fourth aspect of the invention, a method for producing an aircraft according to the second aspect of the invention is provided, comprising the following steps: - positioning the center of mass of the aircraft in such a way that one or more of the drive devices generate a specific predetermined thrust vector assigned to them in each case, so that the aircraft performs a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear when - the associated axes of rotation of two of the at least three drive devices are aligned essentially in the first direction, and the associated axis of rotation of a further one of the at least three drive devices is aligned essentially in the second direction, and - each of the two propulsion devices with axes of rotation aligned in the first direction during hovering rotates substantially in the same direction of rotation about the respectively associated axis of rotation.
In the preferred case that the aircraft comprises at least four drive devices, the associated axes of rotation of two of the at least four drive devices are aligned substantially in the first direction, and the associated axes of rotation of two further ones of the at least four drive devices are aligned substantially in the second direction, and each of the two drive devices with axes of rotation aligned in the first direction during hovering rotates substantially in the same direction of rotation about the respectively associated axis of rotation, and/or each of the two drive devices with axes of rotation aligned in the second direction during hovering rotates substantially in the same direction of rotation about the respectively associated axis of rotation. According to a fifth aspect of the invention, a method is provided for controlling an aircraft with an aircraft body which defines a longitudinal direction, a vertical direction and a transverse direction, the longitudinal direction corresponding to the direction from the tail to the nose of the aircraft, the vertical direction corresponding to the direction of gravity when the aircraft is resting on the ground, and the transverse direction being perpendicular to the longitudinal direction and the vertical direction, and at least two drive devices which are rotatable about a respectively associated axis of rotation in order to generate a respectively associated thrust vector, a first number of drive devices being arranged along a first straight line which runs parallel to the transverse direction, and a second number of the drive devices being arranged along a second straight line which runs parallel to the transverse direction, the first straight line being spaced from the second straight line, and the center of mass of the aircraft being positioned between the first straight line and the second straight line with respect to the longitudinal direction.
The method comprises the following steps: - determining the associated thrust vectors such that the aircraft performs a hover flight when each of the axes of rotation associated with the at least two drive devices is aligned substantially in the transverse direction of the aircraft body, and each of the at least two drive devices rotates substantially in the same direction of rotation about the respectively associated axis of rotation, wherein, in hovering flight, all forces acting on the aircraft and all torques acting on the aircraft with respect to the centre of mass of the aircraft essentially disappear, - driving each of the drive devices essentially in the same direction of rotation such that the respective drive device generates the specific associated thrust vector. According to a sixth aspect of the invention, a method is provided for controlling an aircraft with an aircraft fuselage and at least three drive devices which are mounted around the aircraft fuselage and which are each rotatable about an associated axis of rotation in order to generate a respective associated thrust vector, which comprises the following steps: - determining the associated thrust vectors in such a way that the aircraft performs a hover flight when two of the axes of rotation associated with the at least three drive devices are aligned essentially in the first direction and rotate essentially in the same direction of rotation about the respective associated axis of rotation, and/or another of the axes of rotation associated with the at least three drive devices is aligned essentially in a second direction which is not parallel to the first direction, wherein in hover flight all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear, - aligning the associated axes of rotation of two of the at least three drive devices essentially in the first direction, and aligning the associated axis of rotation of the other of the at least three drive devices substantially in the second direction, - driving each of the drive devices such that the respective drive device rotates in an associated direction of rotation and generates the particular associated thrust vector, wherein each of the drive devices with axes of rotation aligned substantially in the first direction rotates substantially in the same direction of rotation about the respectively associated axis of rotation.
Preferably, the method is provided for controlling an aircraft with at least four drive devices and comprises the following steps: - determining the associated thrust vectors such that the aircraft performs a hover flight when two of the axes of rotation assigned to the at least four drive devices are aligned essentially in a first direction and rotate essentially in the same direction of rotation about the respectively assigned axis of rotation, and/or two further axes of rotation assigned to the at least four drive devices are aligned essentially in a second direction that is not parallel to the first direction and rotate essentially in the same direction of rotation about the respectively assigned axis of rotation, wherein in hover flight all forces acting on the aircraft and all torques acting on the aircraft with respect to the center of mass of the aircraft essentially disappear, - aligning the associated axes of rotation of two of the at least four drive devices essentially in the first direction, and aligning the associated axes of rotation of the two further ones of the at least four drive devices essentially in the second direction, - driving each of the drive devices such that the respective drive device rotates in an associated direction of rotation and generates the determined associated thrust vector, wherein each of the drive devices with axes of rotation aligned substantially in the first direction rotates substantially in the same direction of rotation about the respectively associated axis of rotation and/or each of the two drive devices with axes of rotation aligned substantially in the second direction rotates substantially in the same direction of rotation about the respectively associated axis of rotation.
Preferably, in the methods for controlling an aircraft according to the fifth or sixth aspect, all of the determined associated thrust vectors are selected to be approximately identical. Advantageously, the methods for controlling an aircraft according to the fifth or sixth aspect further comprise the following step: - positioning the center of mass of the aircraft in such a way that all forces acting on the aircraft and all forces relating to the The torques acting on the aircraft due to the centre of mass of the aircraft essentially disappear when the propulsion devices generate the specific predetermined thrust vector assigned to them. The advantages of the methods according to the third to sixth aspects of the invention are the same as those already described in connection with the aircraft according to the invention according to the first and second aspects. The expedient, advantageous and preferred embodiments of the first and second aspects therefore apply accordingly to the third to sixth aspects of the invention.
Preferably, in the aircraft or method according to any of the aspects of the invention, each of the drive devices is structurally identical. Particularly preferably, for any aircraft or method according to any of the aspects of the invention, the propulsion devices comprise cyclogyro rotors. Each cyclogyro rotor expediently comprises a plurality of rotor blades which can be rotated along a circular path about the respectively associated axis of rotation of the drive device or of the cyclogyro rotor; a pitch mechanism with a coupling device and a bearing device, wherein each of the plurality of rotor blades is pivotally mounted by the bearing device about its rotor blade bearing axis parallel to the axis of rotation of the drive device or of the cyclogyro rotor. Furthermore, the cyclogyro rotor expediently comprises an offset device to which each rotor blade is coupled by the coupling device at a connection point assigned to it.
The offset device defines an eccentric bearing axis that is mounted at an adjustable offset distance parallel to the axis of rotation of the drive device or the cyclogyro rotor. As a result, the rotation of the rotor blades along the circular path around the axis of rotation of the drive device or the cyclogyro rotor causes a pitch movement of the rotor blades if the offset distance is set to a value other than zero. In general, however, the requirement for the lift force of an aircraft is largely constant, and an increase is usually not needed, since the main purpose here is to counteract gravity. With the help of the offset device, however, the thrust force can be reduced due to the increase, which results in a reduced power consumption of the rotor. In the following, preferred embodiments of the present invention are described with reference to the following figures. They show: Figure 1: a perspective view of an aircraft according to the first aspect of the invention; Figure 2a: a schematic representation of a drive device and the forces and torques acting on it; Figure 2b: a schematic representation of a drive device in forward flight of the aircraft and the forces and torques acting on it, taking into account an incoming flow; Figure 3a: a schematic representation of the aircraft according to the first aspect of the invention in plan view; Figure 3b: a schematic representation of the aircraft according to the first aspect of the invention and the forces and torques acting on it in side view; Figure 3c: an example configuration of an aircraft with four parallel and equally sized drive devices to illustrate the preferred center of mass position of the aircraft; Figure 4: a schematic representation of the aircraft according to the first aspect of the invention in plan view to generalize the conditions for a stable flight position; Figure 5: a perspective view of a drive device according to the invention; Figure 6: a perspective view of an aircraft according to the second aspect of the invention;
Figure 7a: a schematic representation of the aircraft according to the second aspect of the invention in plan view and of the forces and torques acting on it; Figure 7b: a schematic representation of the aircraft in a configuration according to the invention according to the second aspect of the invention and of the forces and torques acting on it in a first side view; Figure 7c: a schematic representation of the aircraft in a configuration according to the invention according to the second aspect of the invention and of the forces and torques acting on it in a second side view; Figure 7d: an example configuration of an aircraft according to the second aspect of the invention with four drive devices arranged in a star shape and of equal size to illustrate the preferred center of mass position of the aircraft; Figure 8a: a section of an aircraft with n drive devices according to the second aspect of the invention in plan view to explain the determination of the center of mass; Figure 8b: a section of the aircraft with n drive devices in side view; Figure 9a: a schematic representation of an aircraft according to the second aspect of the invention with three drive devices; Figure 9b: a schematic representation of an aircraft according to the second aspect of the invention with seven drive devices; Figure 9c: a schematic representation of an aircraft according to the second aspect of the invention with six drive devices.
Figure 1 shows a perspective view of an aircraft 100 according to the first aspect of the invention with an aircraft body 120 and several drive devices 1F, 1R. Each the drive devices 1F, 1R can be mounted on the aircraft body 120 using appropriate mounting or storage devices. The illustrated aircraft 100 can be, for example, an aircraft, a manned aircraft, a drone or a so-called Micro Air Vehicles (MAVs) trade. To further describe the aircraft, a coordinate system is introduced which has a longitudinal direction 101 or 102. Longitudinal axis, a transverse direction 102 or Transverse axis and a vertical direction 103 respectively. vertical axis defined. The coordinate system should be firmly anchored to the aircraft 100. The reference directions 101, 102, 103 or axes are defined as follows: The longitudinal direction 101 corresponds to the direction from the tail 122 to the nose 121 of the aircraft 100. In the embodiment shown in Fig.1, the longitudinal direction 101 is thus in a horizontal plane (parallel to the ground when the aircraft 100 is resting on the ground), and extends from the tail 122 (i.e. the rear part) of the aircraft 100 to the bow 121, or nose 121, (i.e. the front part) of the aircraft 100.
The vertical direction 103 or axis corresponds to the direction of gravity when the aircraft 100 rests on the (flat) ground. In other words: the vertical direction 103 is perpendicular to the above-mentioned horizontal plane, which includes the longitudinal direction 101. The transverse direction 102 or axis is perpendicular to both the longitudinal direction 101 and the vertical direction 103. In other words: the transverse direction 102 lies in the above-mentioned horizontal plane, which includes the longitudinal direction 101, and is perpendicular to the longitudinal direction 101. The aircraft 100 shown has four propulsion devices 1F, 1R. The drive devices 1F, 1R shown are cyclogyro rotors. The aircraft 100 shown in Fig. 1 can therefore also be referred to as a cyclogyro. The drive devices are described in more detail in connection with Figure 5. Each of these drive devices 1F, 1R is mounted so as to be rotatable about an axis of rotation 5 assigned to it. Each drive device 1F, 1R comprises several rotor blades 2 which are pivotably mounted about their longitudinal axis.
This allows the angle of inclination of the rotor blades 2 to be varied during the rotation of the drive device 1F, 1R. By controlling the rotation speed (hereinafter also referred to as rotation speed) of the drive devices 1F, 1R and controlling the inclination angle of the rotor blades 2 the magnitude and direction of the generated thrust force or the thrust vector describing it can be varied. In Fig.1 it can be seen that two of the four drive devices 1F are arranged in the front (bow) area of the aircraft 100, two further drive devices 1R in the rear (tail) area of the aircraft 100. The front and rear areas of the aircraft are defined as follows: The total length of the aircraft is measured in the longitudinal direction 101; the frontmost part of the aircraft (i.e. the nose 121 of the aircraft 100) is assigned the relative longitudinal coordinate 0, the rearmost part 122 of the aircraft 100 is assigned the relative longitudinal coordinate 100%. In this convention, the front part or
Area is defined as corresponding to the (longitudinal) range from 0 to 40% of the total length of the aircraft, the rear part or Range that corresponds to the (longitudinal) range from 60% to 100% of the total length of the aircraft. The two drive devices 1F in the front area lie on a common straight line that runs parallel to the transverse direction 102 or axis; likewise, the two drive devices 1R in the rear area lie on a common straight line that runs parallel to the transverse direction 102 or axis. It should be noted that the straight lines mentioned do not necessarily have to be a common axis of rotation to which the drive devices are (rigidly) coupled. Each drive device 1F, 1R can rotate via its own axis of rotation 5 assigned to it, and it is also possible for each of the drive devices 1 to be controlled individually, in particular to control its rotational speed separately. Furthermore, according to the invention it is not necessary for all drive devices 1F, 1R to be located in the same horizontal plane.
As shown in Fig. 1, it may be expedient if the two drive devices 1R in the rear area of the aircraft are arranged higher than the two drive devices 1F in the front area. This has the advantage that the propulsion devices 1R in the rear area receive a better airflow and are less affected by the air eddies and turbulence caused by the propulsion devices 1F in the front area. In the embodiment of Fig.1, the rotation axes 5 assigned to the drive devices 1F, 1R are aligned parallel to the transverse direction 102. According to the invention, however, it is not absolutely necessary that all axes of rotation 5 are parallel to each other get lost. It is already according to the invention if each of the associated axes of rotation 5 is aligned substantially in the transverse direction 102 of the aircraft body 120. According to the invention, an axis of rotation 5 is aligned substantially in the transverse direction 102 of the aircraft body 120 if the angle enclosed between the axis of rotation 5 and an axis which runs in the transverse direction and intersects the axis of rotation 5 is less than 45°, preferably less than 30°, particularly preferably less than 15°.
The term “substantially aligned in the transverse direction” does not therefore exclude the possibility that the axes of rotation 5 are also exactly parallel to one another. The aircraft 100 according to the invention is designed such that it can perform a hover flight by rotating each of the four drive devices 1F, 1R shown in the same direction of rotation about the respectively associated axis of rotation 5. The resulting design limitations for the aircraft 100 are explained in connection with the other figures, in particular figures 3a and 3b. In the generalized case that the axes of rotation 5 are aligned substantially in the transverse direction 102 of the aircraft body 120, the invention requires that each of the drive devices 1 rotates substantially in the same direction of rotation about the axis of rotation 5 assigned to it. As already explained in detail in the introduction, this is fulfilled if the scalar product of the vector of the angular velocity of a specific drive device 1F, 1R and a fixed but arbitrary vector pointing in the transverse direction 102 has the same sign for all drive devices 1R, 1F.
Figure 2a illustrates the force 7 and the torque 8 acting on a drive device 1 rotating at a certain rotational speed about a rotation axis 5. In Fig. 2a only the front view of the drive device 1 is shown, schematically. In the case shown, it is assumed that no air flows through the drive device 1. In the case shown, the drive device 1 rotates clockwise. The angular velocity vector corresponding to this rotation thus points into the plane of the paper (according to the right-hand rule). The thrust vector F, 7 acting on the drive device 1 is perpendicular to the axis of rotation 5 of the drive device 1. If cyclogyro rotors are used as propulsion devices 1, the thrust vector F, 7 is determined by the periodic adjustment of the rotor blades of the Cyclogyro rotors are produced. With the help of an offset device of the cyclogyro rotor, the periodic rotor blade adjustment can be changed and thus the thrust vector can be rotated in the entire plane which is normal to the axis of rotation 5 of the cyclogyro rotor and the amount of the thrust vector can be changed.
For this purpose, a thrust vector control is conveniently used. In addition to the thrust vector F, 7, the drive device 1 generates a torque M, 8 about the axis of rotation 5 against the direction of rotation 51. This torque M, 8 about the axis of rotation 5 results from the air forces (lift and drag forces), or their tangential components, of the drive device 1; in the case of a cyclogyro rotor, the air forces are primarily due to the rotating rotor blades. In order to maintain a constant rotational speed, the drive device 1 must therefore generate a (drive) torque that counteracts the torque resulting from the air forces. However, in order for the drive device 1 to be able to generate such a (drive) torque also during the flight phase, a further torque M, 8 is required, which the aircraft body must apply (according to the principle actio = reactio) in order to “support” the drive device 1 in the air. In order to maintain a constant rotational speed against the air forces, this latter torque M, 8 is (neglecting dissipative effects) approximately equal in magnitude to the torque generated by the air forces, and also points in the same direction as the latter.
Since the torque generated by the air forces counteracts the direction of rotation 51 of the drive device 1, the torque M, 8 applied by the aircraft body also counteracts the direction of rotation 51 of the drive device 1. Assuming that the torque due to the air forces and that of the propulsion device are essentially equal in magnitude but oppositely directed, the torque M, 8 applied by the aircraft body remains as the net torque due to the rotation of the propulsion device 1. This torque M, 8 is therefore equivalent to the drive torque of the drive device 1. The torque M, 8 can therefore be directly related to the magnitude of the thrust vector F, 7. The design limitations of the aircraft according to the invention already mentioned in connection with Figure 1 and further described with regard to Figures 3a and 3b can therefore be overcome by using a mathematical-physical relation between the torque M, 8 and the thrust vector F, 7.
Mathematically (and physically) the relationship between the thrust force or corresponding thrust vector F, 7 and the (drive) torque M, 8 can be explained based on general equations of a propeller. Due to the position of the rotor blades in relation to the axis of rotation, a classic propeller differs from a cyclogyro rotor, but in both concepts the thrust generation is based on the targeted displacement of air in one direction by the rotor blades. The equations used below are derived in the appendix to this description for the sake of completeness. First, we consider the power required to displace the air. This performance P<sub>Luft</sub>can be obtained from the so-called beam theory (see Appendix), which leads to the following expression: (1) <img file="imgf000032_0001.tif" frnum="0001" he="10" id="imgf000032_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0011" wi="35"/> where F is the magnitude of the thrust vector and V<sub>a</sub>is the total flow velocity of the air in the plane of the propulsion device.
The plane of the drive device is a plane that passes through the axis of rotation of the drive device and is perpendicular to the direction of air flow and thus to the thrust vector F. This power is provided via the drive device 1. First, the general power P<sub>Antrieb</sub>of the drive device is: <img file="imgf000032_0002.tif" frnum="0001" he="9" id="imgf000032_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0012" wi="39"/> (2) where M is the magnitude of the (drive) torque M, 8 and ω is the rotational speed (magnitude of the angular velocity vector) of the drive device 1. The relationship between the two powers P<sub>Luft</sub>and P<sub>Antrieb</sub>can be described via the efficiency η as follows: (3) <img file="imgf000032_0003.tif" frnum="0001" he="10" id="imgf000032_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0013" wi="42"/> The efficiency η indicates how effectively the drive power P<sub>Antrieb</sub>is converted into an air flow.
The ratio between the rotational speed ω and radius r, 52 of the drive device 1, on the one hand, and the total flow velocity V<sub>a</sub>, on the other hand, is a dimensionless characteristic of the drive device 1 and is denoted here by H (in the case of propellers this is usually called "progression degree"): (4) <img file="imgf000033_0001.tif" frnum="0001" he="13" id="imgf000033_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0014" wi="28"/> The relationship between the (drive) torque M, 8 and the thrust force or the thrust vector F, 7 can then be established starting from equation (3) and inserting the formulas (1), (2) and (4). (5) <img file="imgf000033_0002.tif" frnum="0001" he="13" id="imgf000033_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0015" wi="32"/> This relationship only depends on the parameters H, r and η of the drive device 1.
The relationship between (the amounts of) (drive) torque M, 8 and thrust or Thrust vector F, 7 can therefore be described as a linear function with a general proportionality factor a: <img file="imgf000033_0003.tif" frnum="0001" he="8" id="imgf000033_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0016" wi="27"/> (6) This relationship will be used later. Figure 2b shows schematically a propulsion device 1 in forward flight. The direction of movement of the aircraft, which comprises the drive device 1 shown, is indicated by the arrow 110. The torque M, 8, which corresponds to the drive torque of the drive device 1, has already been described in connection with Figure 2a. It is shown that the drive device 1 is supplied with air from the outside 9. The inflow 9 of air changes the aerodynamic properties of the propulsion device 1 and thus the properties of the generated thrust vector.
If the aircraft and thus the propulsion device 1 is in forward flight, the propulsion device 1 is actively supplied with air from the front. As already explained in the introduction, the changed properties of the drive device 1 can be approximately explained by the Magnus effect, which states that a rotating round body in a flow experiences a transverse force normal to the flow direction. The direction of the transverse force depends on the direction of rotation 51 of the body, here, the drive device 1. Due to the Magnus effect, in addition to the thrust force described with respect to Figure 2a, the vertical component of which is designated F<sub>rotor</sub>, 71 in Figure 2b, an additional thrust force or an additional contribution F<sub>magnus</sub>, 72 to the thrust vector in the vertical direction is generated. This means that the entire thrust force acting in the vertical direction, the so-called buoyancy force of the propulsion device 1 is increased. In general, however, the requirement for the lift force of an aircraft is largely constant and an increase is usually not needed, since the main purpose here is to counteract the force of gravity.
Due to the noticeable contribution F<sub>magnus</sub>, 72 to the thrust vector during forward flight, the contribution F<sub>rotor</sub>, 71 of the thrust vector generated by the propulsion device 1 can be reduced. This is associated with a reduced power consumption of the drive device 1. Put simply, the Magnus effect replaces part of the thrust of the propulsion device 1 and thus reduces the power requirement in forward flight compared to hovering flight. However, if the propulsion device 1 were to rotate in the opposite direction with a constant flow 9, the additional transverse force F<sub>magnus</sub>, 72 of the Magnus effect would act against the thrust force F<sub>rotor</sub>, 71 and thus reduce the total thrust force or increase the power requirement while the same lift force is required. In the aircraft according to the invention, the described positive effect of the Magnus effect is exploited in that during hovering and forward flight of the aircraft all drive devices rotate in the same direction about the associated axes of rotation.
In a generalized arrangement with axes of rotation aligned substantially in the transverse direction of the aircraft body, the drive devices rotate substantially in the same direction of rotation, as explained in more detail above. If the propulsion devices 1 rotate in essentially the same direction of rotation about the respective associated axis of rotation, the contribution to the lift force by the transverse force F<sub>magnus</sub>, 72 becomes greater the faster the aircraft flies in forward flight. This means that it is sufficient to configure the aircraft in hover mode, where the The air flow velocity 9 is usually at its lowest in order to ensure a stable flight attitude of the aircraft during forward flight. The conditions for a stable flight attitude in hovering flight and in forward flight (equilibrium of all forces and torques acting on the aircraft) have already been generally stated in the introduction; in the following, in connection with Figures 3a and 3b, design restrictions for the aircraft according to the first aspect of the inventions are derived from these conditions.
In Figure 3a, an aircraft 100 according to the first aspect of the invention is shown in a highly schematic representation in plan view. In addition to the aircraft body 120 already described in connection with Figure 1, the drive devices 1F and 1R, the axes of rotation 5 and the longitudinal direction 101 and transverse direction 102 assigned to them, the center of mass S, 150 of the aircraft 100 can also be seen. The location or Positioning of the center of mass S, 150 is of central importance for compensating the equally directed torques caused by the drive devices 1 rotating in essentially the same direction of rotation. This is described in more detail with regard to Figure 3b. Figure 3b shows the aircraft shown in plan view in Figure 3a according to the first aspect of the invention in side view and in a highly schematic representation. In this side view, only one of the two propulsion devices 1F arranged in the front area of the aircraft and one of the two propulsion devices 1R arranged in the rear area of the aircraft can be seen.
Furthermore, in Fig. 3b, the four drive devices 1F and 1R are arranged in a horizontal plane. However, the following statements also apply in the event that not all drive devices are in a horizontal plane. The axes of rotation associated with the drive devices 1F and 1R are parallel to each other and parallel to the transverse direction (which points into the plane of the sheet). According to the invention, all four drive devices 1F, 1R rotate in the same direction of rotation 51 with a certain associated rotational speed. In Fig. 3b, all drive devices 1F and 1R rotate clockwise, which means that all four drive devices are clockwise with respect to the transverse direction (y-axis) indicated in Fig. 3a. right-handed. In other words, the scalar product of each of the angular velocity vectors associated with the drive devices 1F, 1R with the unit vector in the transverse direction is positive. Regardless of the reference system used, one can also say that the propulsion devices rotate in such a way that the surface of the propulsion devices that first encounters the air flow during forward flight rotates against the direction of gravity.
When the drive devices rotate clockwise, the Magnus effect has a particularly positive effect. This applies to any number of drive devices. As already mentioned above, a thrust vector is generated by the rotation of each drive device 1F, 1R. In the notation according to Fig. 3b, the thrust vector jointly generated by the two drive devices 1F arranged in the front area is designated F<sub>1</sub>, 701, and the thrust vector jointly generated by the two drive devices 1R arranged in the rear area is designated F<sub>2</sub>, 702. Because all drive devices 1F and 1R rotate in the same direction of rotation 51, all resulting (drive) torques M<sub>1</sub>, 81 M<sub>2</sub>, 82 also act in the same direction, where M<sub>1</sub>, 81 denotes the (drive) torque of both front drive devices 1F, and M<sub>2</sub>, 82 denotes the (drive) torque of both rear drive devices 1R. Now the momentum and spin laws are set up around the center of mass S, 150 of the aircraft, whereby in the case shown only the momentum law in the vertical direction 103 (z-axis) and the spin law around the transverse direction (y-axis) are relevant, since only here are forces or
torques act. The conditions for a stable hover are then: (7) (<sub>8)</sub> <img file="imgf000036_0001.tif" frnum="0001" he="16" id="imgf000036_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0017" wi="85"/> The thrust vectors F<sub>1</sub>and F<sub>2</sub>can be adjusted to satisfy the two equilibrium conditions. The thrust vectors are conveniently set by the thrust vector control. l<sub>1</sub>, 131 and l<sub>2</sub>, 132 indicate, in relation to the longitudinal direction, the distance of the center of gravity S, 150 from the drive devices 1F in the front area or 1R in the rear area. F<sub>S</sub>, 160 denotes the weight of the entire aircraft. However, it is also possible to use the two equilibrium conditions to determine the center of mass of the aircraft in such a way that the said conditions for hovering are met for certain given thrust vectors F<sub>1</sub>and F<sub>2</sub>.
The torques M<sub>1</sub>, 81 and M<sub>2</sub>, 82 shown in Fig.3b correspond to the drive torques of the two drive devices 1F and the two drive devices 1R respectively. There is a difference between the magnitudes of the torques M<sub>1</sub>, 81 and M<sub>2</sub>, 82 and the magnitudes of the thrust vectors F<sub>1</sub>, 701 and F<sub>2</sub>, 702. F<sub>2</sub>, 702 of the corresponding drive devices 1F or1R there is a mathematical-physical connection. This is determined by equation (6) given above. This means that the magnitudes of the torques M<sub>1</sub>, 81 and M<sub>2</sub>, 82 are proportional to the generated magnitudes of the thrust vectors F<sub>1</sub>, 701 and F<sub>2</sub>, 711 respectively. F<sub>2</sub>, 702. The torques cannot therefore be freely controlled. As explained above in connection with equation (6), the proportionality factor α of each drive device is essentially dependent on the efficiency of the drive device, its angular velocity and other characteristics of the drive device. Each drive device can have a different proportionality factor α. However, the values of α of different drive devices of the same design or Size typically the same order of magnitude. In terms of functionality, they are essentially identical. According to equation (6), the amounts M<sub>1</sub>, M<sub>2</sub>of the torques M<sub>1</sub>, 81 and
M<sub>2</sub>, 82 can be written as <img file="imgf000037_0001.tif" frnum="0001" he="10" id="imgf000037_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0018" wi="55"/> This gives the torque equation (8) as <img file="imgf000037_0002.tif" frnum="0001" he="10" id="imgf000037_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0019" wi="102"/> This equation can now be converted into a ratio of the amounts F<sub>1</sub>and F<sub>2</sub>of the two thrust vectors F<sub>1</sub>, 701 and F<sub>2</sub>, 702 can be transformed: (9) <img file="imgf000038_0001.tif" frnum="0001" he="17" id="imgf000038_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0020" wi="41"/> Equation (9) can serve as a configuration formula for the aircraft.
Equation (9) initially contains three freely selectable quantities (from the set of F<sub>1</sub>, F<sub>2</sub>, l<sub>1</sub>, l<sub>2</sub>), but in a stable flight attitude equation (7) must also be taken into account, which is why only two of the four quantities mentioned above can be freely selected. There are therefore several possibilities to satisfy equations (7) and (9). (i) In a first case, it can be required that the aircraft is designed symmetrically. This means that the front axes of rotation 5, i.e. the axes of rotation of the drive devices 1F arranged in the front area of the aircraft, and the rear axes of rotation 5, i.e. the axes of rotation of the drive devices 1R arranged in the rear area of the aircraft, are equidistant from the center of mass S, 150. In other words, the centre of mass S, 150 is located in the middle between the front and rear axes of rotation 5 with respect to the longitudinal direction.
In this case l<sub>1</sub>= l<sub>2</sub>. Then, from equation (9) and because the front <img file="imgf000038_0002.tif" frnum="0001" he="11" id="imgf000038_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0021" wi="21"/> propulsion devices 1F must produce more thrust than the rear propulsion devices 1R, so F<sub>1</sub>> F<sub>2</sub>. Thus, the front drive devices 1F must be designed larger than the rear drive devices 1R. In this case, the center of mass S, 150 will therefore tend to move forward, which means that l<sub>1</sub>< l<sub>2</sub>, and the required thrust vectors F<sub>1</sub>and F<sub>2</sub>of the drive devices 1F and 1R increase further. (ii) In a second case, the drive devices 1F and 1R are particularly preferably designed to be structurally identical.
This means that they are identical in construction and, for example, have the same size, the same wingspan, the same number of rotor blades, the same diameter and/or generate similar or identical (maximum) thrust forces/thrust vectors. In this case, F<sub>1</sub>= F<sub>2</sub>or F<sub>1</sub>§ F<sub>2</sub>. With F<sub>1</sub>= F<sub>2</sub>Ł F, equation (7) initially yields F = F<sub>S</sub>/2. From equation (9) we then obtain <img file="imgf000038_0003.tif" frnum="0001" he="8" id="imgf000038_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0022" wi="37"/> If the distance in the longitudinal direction between the front drive devices 1F and rear drive devices 1R is l = l<sub>1</sub>+ l<sub>2</sub>, then the last equation gives: (10) (11) <img file="imgf000039_0001.tif" frnum="0001" he="24" id="imgf000039_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0023" wi="50"/> It can be seen that the center of mass S, 150 of the aircraft is shifted in the longitudinal direction from the center l/2 between the front axes of rotation 5 of the front drive devices 1F and the rear axes of rotation 5 of the rear drive devices 1R in the direction of the rear axes of rotation 5 of the rear drive devices 1R, namely by (a<sub>1</sub>+ a<sub>2</sub>)/2.
Typically in this case a<sub>1</sub>= a<sub>2</sub>≡ a. If an aircraft is now built with structurally identical, equally sized propulsion devices 1F and 1R and thus approximately equally large thrust forces / thrust vectors F<sub>1</sub>, 701 or F<sub>2</sub>, 702 is configured per pair of drive devices 1F or 1R, the center of mass S, 150 can therefore be optimally positioned so that the torques M<sub>1</sub>, 81 or M<sub>2</sub>, 82 can be compensated purely by the position of the center of mass S, 150. The said optimal position is determined by equations (10) and (11). Here and in the following, it must be noted that only the position of the drive devices and the center of gravity in the longitudinal direction 101 plays a role in the considerations. The storage or Positioning of drive devices and center of mass with respect to the transverse direction and vertical direction 103 is not relevant here and is at the discretion of the expert.
A storage or arrangement that is as symmetrical as possible. However, positioning in the latter two directions is preferable. (iii) According to the invention, it is also possible that aspects of the first case (i) and the second case (ii) are combined with each other. This means that the center of mass S, 150 of the aircraft can be located between the front and rear axes of rotation of the propulsion devices 1F and 1B respectively. 1R can be shifted in such a way that the conditions (7) and (8) for a stable hover at certain given, also different thrust vectors / thrust forces of individual drive devices are fulfilled. For practical applications, it is not always possible to place the masses in an aircraft such that the total center of mass S, 150 can be positioned exactly at the predetermined optimal position described in case designs (i), (ii) or (iii); for example, for case design (i) l<sub>1</sub>= l<sub>2</sub>; for case design (ii) l<sub>1</sub>and l<sub>2</sub>are given by equations (10) and (11).
Therefore, an area is defined below in which the center of mass S, 150 can lie, so that it is still possible to compensate the torque with the thrust forces / thrust vectors F<sub>1</sub>, 701 or. F<sub>2</sub>, 702 of the pairs of drive devices 1F respectively. to support 1R. For this purpose, it is assumed that a pair i of drive devices has a maximum permissible (i.d. R. predetermined) thrust / a maximum permissible thrust vector of F<sub>i,max</sub>. It is assumed that F<sub>i,max</sub>is greater than or equal to the thrust forces F<sub>i,opt</sub>corresponding to the optimal configuration. This is because an aircraft requires at least the thrust forces F<sub>i,opt</sub> to remain in a stable hover; in the preferred case, each pair of propulsion devices still provides an excess of thrust, which can be used, among other things, to deviate the position of the center of gravity S, 150 from the optimal position.
F<sub>i,max</sub>is the maximum thrust force of a drive device permitted by the thrust vector control, which must therefore always be greater than or equal to the thrust force for the optimal design F<sub>i,opt</sub>. Taking into account the principle of momentum according to equation (7), the following applies: <img file="imgf000040_0004.tif" frnum="0001" he="11" id="imgf000040_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0024" wi="106"/> This allows a maximum permissible thrust vector ratio to be defined: <img file="imgf000040_0003.tif" frnum="0001" he="17" id="imgf000040_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0025" wi="50"/> And correspondingly: <img file="imgf000040_0002.tif" frnum="0001" he="10" id="imgf000040_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0026" wi="102"/> and thus a minimum permissible thrust vector ratio of <img file="imgf000040_0001.tif" frnum="0001" he="18" id="imgf000040_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0027" wi="50"/> These thrust vector ratios F<sub>1</sub>/F<sub>2</sub> are also described by equation (9); using the latter, the maximum permissible distance in the longitudinal direction of the center of mass S, 150 from the front axes of rotation 5 to <img file="imgf000041_0002.tif" frnum="0001" he="13" id="imgf000041_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0028" wi="68"/> and the minimum permissible distance in the longitudinal direction of the center of mass S, 150 from the front axes of rotation 5 can be calculated.
<img file="imgf000041_0001.tif" frnum="0001" he="18" id="imgf000041_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0029" wi="69"/> If the centre of mass S, 150 is outside the range (12) <img file="imgf000041_0003.tif" frnum="0001" he="8" id="imgf000041_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0030" wi="51"/> it is no longer possible to compensate for the deviation of the centre of mass S, 150 from the optimum position according to equation (10) by the thrust forces F<sub>1</sub>, 701 or. F<sub>2</sub>, 702 of the drive devices 1F and 1R respectively. Figure 3c serves to illustrate the region described above in which the center of mass S, 150 of the aircraft can expediently be located for implementing the invention according to the first aspect.
Fig.3c shows schematically an aircraft with propulsion devices 1F, 1R, which are arranged along two straight lines, each running parallel to the transverse direction of the aircraft. Conveniently, the aircraft comprises four propulsion devices 1F, 1R, of which two 1F are arranged in the front area and two 1R in the rear area, as already described in connection with Figures 3a and 3b. It is further assumed that the drive devices 1F, 1R are structurally identical (as in case (ii)), here in particular: a<sub>1</sub>= a<sub>2</sub>≡ a. First, it is further assumed that the torque compensation is to be realized purely via the position of the center of mass S, 150, whereby <img file="imgf000041_0004.tif" frnum="0001" he="7" id="imgf000041_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0031" wi="99"/> <sub>, , , ,</sub>applies.
For the aircraft embodiment considered here, a total weight force of F<sub>s</sub>= 1000 N generated by a corresponding total mass is assumed; the characteristic number / proportionality factor is typically a = 0.2 m; the distance between the drive devices in the longitudinal direction is defined as. <img file="imgf000042_0002.tif" frnum="0001" he="14" id="imgf000042_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0032" wi="49"/> Based on these specifications, equations (10) and (11) result in an optimal center of mass position of <img file="imgf000042_0001.tif" frnum="0001" he="38" id="imgf000042_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0033" wi="74"/> If it is not possible to place the total center of mass S, 150 of the aircraft at the position l<sub>1,opt</sub>= 1.2 m, a range is now defined in which the position of the center of mass S, 150 can be located, so that the torque compensation can be compensated by the thrust forces / thrust vectors of the drive devices 1F, 1R.
For this purpose, the maximum permissible thrust that can be generated by all propulsion devices arranged along a straight line, which is conveniently controlled by the thrust vector control, is defined as <img file="imgf000042_0003.tif" frnum="0001" he="8" id="imgf000042_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0034" wi="36"/>. This specification allows the maximum and minimum permissible thrust vector ratio <img file="imgf000042_0004.tif" frnum="0001" he="12" id="imgf000042_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0035" wi="67"/> and the range for the position of the center of mass according to equation (12) <img file="imgf000042_0005.tif" frnum="0001" he="15" id="imgf000042_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0036" wi="72"/> to be calculated.
That is, in this example, the longitudinal centre of gravity is conveniently located 1.1 to 1.3 m from the front axes of rotation of the corresponding front drive devices 1F. Figure 4 shows a further embodiment of an aircraft 100 according to the first aspect of the invention. This Fig. 4 serves primarily to generalize the results derived in connection with Figures 3a, 3b and 3c for any number K > 2 of drive devices 1. It has already been pointed out above that for the In the context of the invention, the positioning of the drive devices 1 in the longitudinal direction is primarily important. The drive devices can be positioned at different heights in the vertical direction. The longitudinal direction is marked as x-axis 101 in Fig. 4. It is assumed that the K propulsion devices of the aircraft are arranged along N > 1 straight lines gi. As already explained above, the said straight lines are not structural components of the aircraft 100, but merely serve to illustrate the geometric arrangement of the propulsion devices 1.
On a certain straight line g<sub>i</sub> (denoted by index i, i = 1, … ,N), n<sub>i</sub>, i = 1, … ,N, drive devices 1 should be arranged. This means that <img file="imgf000043_0001.tif" frnum="0001" he="12" id="imgf000043_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0037" wi="33"/> Furthermore, it is assumed that all n<sub>i</sub> drive devices 1 arranged on a straight line g<sub>i</sub> with index i generate a total thrust force / a total thrust vector with magnitude (where F<sub>ij</sub> is the thrust vector generated by the j-th drive device arranged on the straight line <img file="imgf000043_0003.tif" frnum="0001" he="9" id="imgf000043_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0038" wi="24"/> g<sub>i</sub>); the magnitude of the total (drive) torque of all n<sub>i</sub> drive devices arranged on the straight line gi with index i is Mi. For each , i = 1, ... ,N, the following relationship therefore applies according to equation (6): <img file="imgf000043_0002.tif" frnum="0001" he="10" id="imgf000043_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0039" wi="34"/> where for each straight line g<sub>i</sub> with index i a key figure / a proportionality factor a<sub>i</sub> is introduced.
It is pointed out that, although the straight lines g<sub>i</sub> along which the drive devices 1 are arranged are aligned parallel to the transverse direction 102, it is not absolutely necessary according to the invention that all the axes of rotation 5 of the drive devices 1 are aligned (mathematically exactly) parallel to one another or to the transverse direction 102. It is sufficient if the axes of rotation 5 of the drive devices 1, especially in hovering flight, are aligned substantially in the transverse direction 102, in the sense defined in the introduction. In Fig.4 it is shown that the rotation axes 5 of some drive devices 1 are not aligned exactly parallel to the transverse direction 102. According to the invention, the drive devices 1 are nevertheless arranged on a straight line g<sub>i</sub> which runs parallel to the transverse direction 102, because their geometric center lies essentially on such a straight line gi,; it is also possible, in order to meet the condition of the arrangement on a parallel straight line, such that the bearing points of the drive devices 1 lie essentially on such a straight line gi.
Each of the straight lines g<sub>i</sub>with index i is located in the longitudinal direction 101 (x-axis) at a point with coordinate x<sub>i</sub>, i = 1, … ,N, where, without loss of generality, x<sub>i</sub>– x<sub>i</sub>– 1 > 0 is assumed. The longitudinal positions xi of the line gi are fixed but arbitrary. The center of mass S, 150 of the aircraft 100 is located at the coordinate X<sub>S</sub> with respect to the longitudinal direction 101. It is pointed out that, whereas in connection with Figures 3a, 3b, 3c the distances l<sub>1</sub>and l<sub>2</sub>of the straight line from the center of mass were considered, here the coordinates relative to the longitudinal direction 101 of the straight line gi are used; this proves to be more appropriate here. Nevertheless, the relationship between the coordinates of the straight line gi and its distances l<sub>i</sub>from the center of mass S, 150 can be easily established: <img file="imgf000044_0001.tif" frnum="0001" he="10" id="imgf000044_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0040" wi="36"/> With the notations introduced, the conditions for a stable levitation or
Forward flight can be generalized from equations (7) and (8) as follows: (13) (14) <img file="imgf000044_0002.tif" frnum="0001" he="24" id="imgf000044_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0041" wi="98"/> Inserting equation (13) yields equation (14): <img file="imgf000044_0003.tif" frnum="0001" he="8" id="imgf000044_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0042" wi="59"/> thus the coordinate X<sub>S</sub>of the center of mass S, 150 is: (<sub>15)</sub> <img file="imgf000044_0004.tif" frnum="0001" he="11" id="imgf000044_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0043" wi="54"/> An intermediate result should be noted here: From equation (15) the coordinate X<sub>S</sub>of the center of mass S, 150 can be calculated if the thrust vectors F<sub>i</sub> are specified; however, equation (13) also provides a further condition that must be fulfilled for a stable flight attitude.
Therefore, not all N thrust vectors F<sub>i</sub>can be specified, but only N – 1. This means that the position X<sub>S</sub>of the center of mass S, 150 for a stable flight attitude, especially hovering, is determined when N – 1 thrust vectors are specified. The values of the given thrust vectors can of course also be the same. The distance in the longitudinal direction of the center of mass S, 150 from the foremost straight line g1 or the propulsion device which, in the longitudinal direction, is closest to the bow 121 or the nose 121 of the aircraft 100, is: <img file="imgf000045_0001.tif" frnum="0001" he="10" id="imgf000045_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0044" wi="70"/> We will now first consider the case in which the propulsion devices 1 arranged on a straight line gi generate approximately equal thrust forces / thrust vectors Fi for each straight line, i.e. The center of mass S, <img file="imgf000045_0004.tif" frnum="0001" he="10" id="imgf000045_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0045" wi="54"/> 150 will therefore optimally be positioned such that the torques Mi generated by the propulsion devices 1 are compensated purely by the position of the center of mass S, 150.
The said optimal position is determined by equations (13) and (15). From equation (13) follows <img file="imgf000045_0002.tif" frnum="0001" he="9" id="imgf000045_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0046" wi="74"/> And thus from equation (15): (16) <img file="imgf000045_0003.tif" frnum="0001" he="12" id="imgf000045_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0047" wi="48"/> In this case it can conveniently be assumed that a<sub>1</sub>≡ a, i = 1, … ,N. A maximum permissible range for the (longitudinal, x-) coordinate X<sub>S</sub>of the center of mass S, 150 can be determined analogously to the considerations for Figure 3b also for the general case described above, using equations (13), (14) and (15).
Figure 5 shows an embodiment of the drive devices that can be used in an aircraft according to the invention. Each of these drive devices 1 is mounted so as to be rotatable about a rotation axis. Each drive device 1 comprises several rotor blades 2, which are pivotably mounted about their longitudinal axis. This allows the angle of inclination of the rotor blades 2 to be varied during the rotation of the drive device 1. By controlling the rotation speed of the drive devices 1 and the control The angle of inclination of the rotor blades 2 allows the magnitude and direction of the generated thrust vector to be varied. Figure 5 shows a perspective view of an embodiment of a drive device 1 according to the invention. The drive device 1 is cylindrical in shape. The drive device 1 shown is a cyclogyro rotor. This drive device 1 comprises five rotor blades 2, each with an associated pitch mechanism 3, an offset device 4 and a disk 11.
Drive systems with a different number of rotor blades are also possible. The rotor blades 2 are mounted so as to be rotatable about a rotation axis of the drive device 1. The offset device 4 defines an eccentric bearing axis which is mounted eccentrically with respect to the axis of rotation of the drive device 1. In Fig. 5, the offset device is shown as an offset disk. The offset disk is freely rotatable around the eccentric bearing axis. The eccentric bearing of the offset disk 4 implies an eccentric bearing of the pitch mechanism 3. The eccentric bearing of the pitch mechanism 3 causes a change in the position of the rotor blades 2 during one revolution about the axis of rotation of the drive device 1. Each of the illustrated pitch mechanisms 3 comprises a coupling device 31 and a bearing device 33. Each rotor blade 2 is pivotally mounted by the corresponding bearing device 33. The rotor blade 2 is mounted about an axis parallel to the axis of rotation of the drive device 1.
This axis is the rotor blade bearing axis 33. The bearing of the rotor blade 2 can be achieved, for example, by means of a bearing means, such as one or more pins, so-called. main pin. The storage means is preferably a part of the storage device 33. The rotor blade bearing axis 33 can pass through the center of mass of the rotor blade 2. Preferably, however, the rotor blade 2 is mounted at a distance from the center of mass. The coupling device 31 of the pitch mechanism 3 couples the rotor blade 2 to the offset device 4 such that the rotor blade 2 performs a pitch movement when it rotates about the axis of rotation of the drive device 1, and under the condition that the eccentric bearing axis does not coincide with the axis of rotation of the drive device 1. An end piece of the coupling device 31 is coupled to the offset device 4 at a connection point. The other end piece of the coupling device 31 is coupled to the rotor blade 2.
The offset disk 4 is freely rotatable. The axis of rotation of the offset disk 4 preferably runs parallel to the axis of rotation of the drive device 1 at a certain offset distance. This results in the eccentric bearing of the offset disk 4 with respect to the axis of rotation of the drive device 1. This offset distance can be adjustable. An offset device 4 with adjustable eccentricity can be realized, for example, by a planetary gear. A pitch movement of the rotor blades 2 occurs when the offset distance is not zero. The coupling device 31 is coupled to the rotor blade 2 at a coupling point 32. For this purpose, the coupling device 31 can comprise a coupling means. In the drive device 1 shown in Fig.5, the coupling device 31 comprises a connecting rod (English “conrod”) and a pin, so-called. Pitch link pin. The pin is a structural embodiment of the coupling means according to the invention.
In the embodiment shown in Fig. 5, the coupling device 31 is coupled to the rotor blade 2 at the coupling point 32 not by direct connection to the rotor blade 2, but by using a connecting element 61. One end of the connecting element 61 is rigidly connected to the rotor blade 2. This connection is preferably made at the rotor blade bearing point. The other end of the connecting element 61 is coupled to the coupling device / connecting rod 31. In this case, the pitch movement is introduced into the rotor blade 2 via the coupling means with the aid of the connecting rod 31 indirectly via the connecting element 61. However, a direct coupling of the coupling device 31 to the rotor blade 2 is also possible according to the invention. Because the coupling device 31 of the pitch mechanism is mounted eccentrically with respect to the axis of rotation of the drive device 1, the coupling point 32 moves on a circular arc relative to the rotor blade bearing axis 33 when the rotor blade 2 rotates about the axis of rotation of the drive device 1.
This causes the pitch movement of rotor blade 2. This is a pendulum movement of the rotor blade 2 around the rotor blade bearing axis 33. The diameter of the drive device 1 corresponds to twice the distance from the axis of rotation of the drive device 1 to the rotor blade bearing axis 33 or point. This diameter is relevant for the wing speed during rotation and therefore relevant for the thrust generated. In exemplary embodiments of the drive device 1 according to the invention, the diameter is in the range between 150 mm and 2000 mm, preferably between 300 mm and 500 mm, particularly preferably it is 350 mm. Furthermore, the drive device 1 shown in Fig. 5 comprises a disk 11. This disk 11 is designed such that it aerodynamically separates the rotor blades 2 from the remaining components of the drive device 1. Such a disk 11 is particularly advantageous in the event that the drive device 1 is operated at higher speeds.
The length of the rotor blades 2 defines the span of the drive device 1. The span of the drive device 1 is the (longitudinal) distance between the two disks 11. The span of one of the cyclogyro rotors that can be used according to the invention is suitably a few centimeters to two meters, preferably between 350 and 420 mm. In the aircraft according to the invention, several cyclogyro rotors are advantageously used. Their ranges should preferably differ by a maximum of 25%, and preferably by a maximum of 10%. Their diameters preferably differ from each other by a maximum of 25%, expediently by a maximum of 10%. The rotor blades 2 shown in Fig. 5 have a symmetrical profile; the invention is not limited to drive devices with rotor blades with a symmetrical profile. The drive device 1 generates thrust or a thrust vector due to two coupled rotational movements. The first rotational movement is the rotation of the rotor blades 2 around the axis of rotation of the drive device 1.
This first rotational movement leads to a movement of the rotor blades 2 along a circular path around the axis of rotation of the drive device. Specifically, the rotor blade bearing axes 33 and 34 move. Rotor blade bearing points along the circular path. Each rotor blade bearing axis 33 is parallel to the longitudinal axis of the rotor blades 2. The longitudinal axis of the rotor blades 2 is parallel to the axis of rotation of the drive device 1. Thus, the longitudinal axis of the rotor blades 2 is also parallel to the rotor blade bearing axis 33. The thrust direction of the drive device 1 is normal to the axis of rotation of the drive device 1. For optimum thrust generation, all rotor blades 2 should be aligned as best as possible to the flow direction at all times. This ensures that each rotor blade 2 makes a maximum contribution to the total thrust. During the rotation of the drive device 1 about its axis of rotation, the inclination of each rotor blade 2 is continuously changed due to the pitch mechanism described above.
Each rotor blade 2 undergoes a periodic change in the angle of inclination or a pendulum movement. This is the pitch movement. The coupling point 32 moves on a circular arc around the rotor blade bearing axis 33. This is the second rotation. The magnitude and direction of the generated thrust force or the associated thrust vector depend on the inclination of the rotor blades 2. Therefore, the distance of the eccentric bearing of the offset device 4 or the pitch mechanism 3 to the axis of rotation of the drive device 1 influences the amount of thrust force / thrust vector generated. By shifting the eccentric bearing of the offset device 4 in the circumferential direction, i.e. at a constant distance from the axis of rotation of the drive device 1, the direction of the generated thrust vector is changed. Although in Fig. 5 pitch mechanisms 3 are shown only on one side of the drive device 1, it may be expedient for stability reasons to also install corresponding pitch mechanisms on the opposite side of the drive device.
The pitch mechanism can also be installed in the middle of the drive device, for example. Figure 6 shows a perspective view of an aircraft 200 according to the second aspect of the invention with an aircraft fuselage 220 and several drive devices 1A and 1B. Four drive devices 1A and 1B can be seen, which are arranged around the aircraft fuselage 220. Each propulsion device 1A and 1B is connected to the aircraft fuselage 220 via an arm 221 or 222. Each of the drive devices 1A and 1B can be mounted on the arms 221 and 222, respectively, with appropriate mounting or bearing devices. The presence of arms 221 or 222 is not essential. The Propulsion devices 1A and 1B may also be coupled to the aircraft fuselage 220 in other ways. The aircraft body 220 and the propulsion devices 1A and 1B are essentially located in one plane. The illustrated aircraft 200 can be, for example, an aircraft, a manned aircraft, a drone or a so-called
Micro Air Vehicles (MAVs) trade. To further describe the aircraft 200 shown, a reference system is introduced which has a first direction 201, a second direction 202 and a vertical direction 203 or. vertical axis defined. The vertical direction 203 or axis corresponds to the direction of gravity when the aircraft 200 is resting on the ground. The vertical direction 203 is perpendicular to the above-mentioned plane in which the aircraft fuselage 220 and the propulsion devices 1A and 1B are located. The first direction 201 and the second direction 202 or the corresponding axes lie in the said plane and are thus perpendicular to the vertical direction. What is essential for the aircraft 200 of the second aspect of the invention considered here is that the first direction 201 and the second direction 202 are not parallel to each other. In the embodiment shown, the first direction 201 and the second direction 202 are perpendicular to each other. The directions thus defined should be firmly anchored to the aircraft 200.
The aircraft 200 shown has four propulsion devices 1A and 1B. The drive devices 1A and 1B shown are cyclogyro rotors. A more detailed description of cyclogyro rotors has already been given in connection with Fig. 5. Each drive device 1A and 1B is mounted so as to be rotatable about an axis of rotation 5 assigned to it. Each drive device 1A and 1B comprises several rotor blades 2 which are pivotably mounted about their longitudinal axis. This allows the angle of inclination of the rotor blades 2 to be adjusted during rotation of the drive device 1A or 1B can be varied. By controlling the rotational speed (hereinafter also referred to as rotational speed) of the drive devices 1A or 1B and by controlling the inclination angle of the rotor blades 2, the amount and direction of the generated thrust force or the thrust vector describing it can be varied. In Fig.6 it can be seen that the four drive devices 1A and 1B essentially form the corners of a rectangle or square.
In the geometric center of this rectangle or The hull is positioned 220 in the square. It is practical for each of the drive devices 1A and 1B to be arranged from the centre or hull equidistant. For this purpose, the arms 221 and 222 can have the same length. In this case, the drive devices 1A and 1B are arranged at the corners of a square. The two drive devices 1A, the opposite corners of the said rectangle respectively. square lie on a common straight line; in the example shown, this straight line is substantially parallel to the first direction 201; likewise, the two drive devices 1B, which are also opposite corners of the said rectangle or square, lie on a common straight line. square, on a common straight line which is substantially parallel to the second direction 202. It should be noted that the straight lines mentioned do not necessarily have to be a common axis of rotation to which the drive devices are (rigidly) coupled. Each drive device 1A, 1B can rotate via its own associated axis of rotation 5A, 5B, and it is also possible for each of the drive devices 1A, 1B to be controlled individually, in particular to control their rotational speed separately.
In the embodiment of Fig.6, the rotation axes 5A assigned to the drive devices 1A are essentially aligned in the first direction 201. In the embodiment of Fig.6, the rotation axes 5B assigned to the drive devices 1B are essentially aligned in the first direction 202. In Fig.6 it can be seen that the axes of rotation 5A, 5B are not aligned exactly parallel to the first direction 201 or the second direction 202. In fact, it is already according to the invention if each of the associated axes of rotation 5A, 5B is substantially aligned in the first direction 201 or second direction 202. According to the invention, an axis of rotation 5A is aligned substantially in the first direction 201 if the angle enclosed between the axis of rotation 5A and an axis running in the first direction 201 and intersecting the axis of rotation 5A is less than 45°, preferably less than 30°, particularly preferably less than 15°. The designation “substantially aligned in the first direction” therefore does not exclude the possibility that the axes of rotation 5A are also exactly parallel to the first direction 201.
The same applies to the rotation axes 5B of the second drive devices 1B and the second direction 202. The aircraft 200 according to the invention is designed such that it can perform a hover flight in that each of the two drive devices 1A shown rotates essentially in the same direction of rotation about the respectively associated axis of rotation 5A, and/or each of the two drive devices 1B shown rotates essentially in the same direction of rotation about the respectively associated axis of rotation 5B. The resulting design limitations for the aircraft 200 are explained in connection with the other figures, in particular figures 7a and 7b. In Figure 7a, an aircraft 200 according to the second aspect of the invention is shown in a highly schematic representation in plan view. Firstly, the aircraft fuselage 220 already described in connection with Figure 6, the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>and 1B<sub>3</sub>, 1B<sub>4</sub>and the axes of rotation 5A and 5B assigned to them can be seen.
5B, the first direction 201 and second direction 202; the first direction 201 is perpendicular to the second direction 202. To describe the mathematical-physical relationships, it is useful to introduce a (Cartesian) orthogonal coordinate system. In Figures 7a and 7b, a Cartesian coordinate system with x, y and z axes is used. It should be noted that, in general, the first and second directions according to the invention do not need to correspond to the axes of a Cartesian coordinate system. The first and second (and possibly further) directions serve to define the axes of rotation of the propulsion devices, while the (Cartesian) orthogonal coordinate system is intended to serve the purpose of mathematically describing the aircraft. In addition, the center of mass S, 250 of the aircraft 200 is shown. The location or Positioning of the center of mass S, 250 is for balancing the forces generated by the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>respectively, rotating in essentially the same direction of rotation.
1B<sub>3</sub>, 1B<sub>4</sub>caused rectified torques of central importance. This is described in more detail with regard to Figure 7b. In the example shown, the center of mass S, 250 is positioned such that the aircraft 200 can exploit the Magnus effect both in forward flight in (positive) first direction 201 (here coinciding with the positive x-direction) and in forward flight in (positive) second direction 202 (here coinciding with the positive y-direction). If the aircraft 200 moves in the Forward flight in the first direction 201, the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> rotate essentially in the same direction of rotation about the associated axes of rotation 5B, advantageously clockwise. As defined above in connection with the first aspect, this means that the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub> are clockwise rotating with respect to the second direction (y-axis) indicated in Fig. 7a.
In other words: The scalar product of each of the angular velocity vectors associated with the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> with the unit vector in the second direction is positive. Independently of the reference system used, one can also say that the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub> rotate in such a way that the surface of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>, which first encounters the air flow during forward flight, rotates against the direction of gravity. If the aircraft 200 moves in the second direction 202 during forward flight, the drive devices 1A<sub>1</sub>, 1A<sub>2</sub> rotate essentially in the same direction of rotation about the associated axes of rotation 5A, advantageously in an anti-clockwise direction. The definition given above applies accordingly. In the coordinate system shown in Fig. 7a, this means that the scalar product of each of the angular velocity vectors assigned to the drive devices 1A<sub>1</sub>, 1A<sub>2</sub> with the unit vector in the first direction is negative.
Regardless of the reference system used, the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub> rotate in such a way that the surface of the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>, which first encounters the air flow during forward flight, rotates against the direction of gravity. Finally, the thrust vectors F<sub>1</sub>, 2001; F<sub>2</sub>, 2002; F<sub>3</sub>, 2003; and F<sub>4</sub>, 2004 are shown, which are generated due to the rotation of the drive devices about the rotation axes 5A and 5B, respectively. 5B can be generated. The thrust vectors F<sub>1</sub>, 2001; F<sub>2</sub>, 2002; F<sub>3</sub>, 2003; and F<sub>4</sub>, 2004 point out of the image plane, which means that lift is generated. In forward flight in the first direction (x-axis) it is also possible that – when the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> rotate in the same direction – the drive devices 1A<sub>1</sub>, 1A<sub>2</sub> rotate in opposite directions, i.e. one clockwise, the other counterclockwise.
The same applies to forward flight in the second direction (y-axis). The direction of the thrust vectors F<sub>1</sub>, 2001; F<sub>2</sub>, 2002; F<sub>3</sub>, 2003; and F<sub>4</sub>, 2004 remains unaffected. Figures 7b and 7c show the aircraft shown in Figure 7a in plan view according to the second aspect of the invention in different side views and in highly schematic representation. In the side view of Fig. 7b, the two drive devices 1A<sub>1</sub>, 1A<sub>2</sub>and one of the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub> can be seen. In the side view of Fig. 7c, the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub>and one of the two drive devices 1A<sub>1</sub>, 1A<sub>2</sub> can be seen. The rotation axes 5A assigned to the drive devices 1A<sub>1</sub>, 1A<sub>2</sub> are parallel to the first direction 201 (here: x-direction); the rotation axes 5B assigned to the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> are parallel to the second direction (here: y-direction) (which points into the sheet plane).
In the considered embodiment according to the invention, the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> are intended to rotate in the same direction of rotation 251 with a certain associated rotation speed. In Fig.7b, the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub> rotate clockwise as defined above. As already mentioned, the rotation of each drive device 1B<sub>3</sub>, 1B<sub>4</sub> generates a thrust vector. In the notation according to Fig. 7b, the thrust vector generated jointly by the two propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub> is denoted by F<sub>34</sub>, 2034, where F<sub>34</sub>= F<sub>3</sub>+ F<sub>4</sub>(cf. Fig.7a). Because the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> rotate in the same direction of rotation 251, all resulting (drive) torques M<sub>34</sub>, 280 also act in the same direction, where M<sub>34</sub>, 280 denotes the (drive) torque of both drive devices 1B<sub>3</sub>, 1B<sub>4</sub>, i.e. M<sub>34</sub>= M<sub>3</sub>+ M<sub>4</sub>.
The drive devices 1A<sub>1</sub>, 1A<sub>2</sub>generate thrust vectors F<sub>1</sub>, 2001; respectively. F<sub>2</sub>, 2002. The direction of rotation of the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub> is not important for the present consideration, which concerns a design of the aircraft that is favorable for forward flight in the primary direction 201. For reasons of symmetry, however, it is preferable to design the aircraft in such a way that a stable flight attitude, in particular a stable forward flight, is possible even with the same rotating propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>. This will be described further below. With regard to Fig.7b, the momentum and angular momentum laws are set up around the center of mass S, 250 of the aircraft, whereby in the case shown only the momentum laws in the vertical direction 203 (z-axis) and the angular momentum laws around the second direction 202 (y-axis) are relevant, since only here are forces or
torques act. The conditions for a stable hover are then:<sub>(17)</sub> (18) <img file="imgf000055_0001.tif" frnum="0001" he="22" id="imgf000055_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0048" wi="104"/> The (magnitudes of) thrust vectors F<sub>1</sub>, F<sub>2</sub>and F<sub>34</sub>can be adjusted to satisfy the two equilibrium conditions. Conveniently, the thrust vectors are adjusted by the thrust vector control. However, it is also possible to use the two equilibrium conditions to determine the center of mass of the aircraft in such a way that the said conditions for hovering are met for certain given thrust vectors F<sub>1</sub>, F<sub>2</sub>and F<sub>34</sub>.
The torque M<sub>34</sub>, 280 shown in Fig. 7b corresponds to the (drive) torque of both drive devices 1B<sub>3</sub>, 1B<sub>4</sub>. As already explained in connection with the first aspect of the invention, there is a mathematical-physical relationship between the magnitude of the torque M<sub>34</sub>, 280 and the magnitude of the thrust vector F<sub>34</sub>. This is determined by equation (6) given above. Each drive device can have a different proportionality factor a. However, the values of a of different drive devices of the same design or Size typically the same order of magnitude. In terms of functionality, they are essentially identical. According to equation (6), the amounts M<sub>1</sub>, M<sub>2</sub>, M<sub>3</sub>, M<sub>4</sub>of the torques can be written as <img file="imgf000055_0004.tif" frnum="0001" he="11" id="imgf000055_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0049" wi="59"/> Since in the embodiment considered, due to the parallel alignment of the axes of rotation of the drive devices 1B<sub>3</sub>, 1B<sub>4</sub>with the same direction of rotation M<sub>3</sub>and M<sub>4</sub>
are parallel, the following also applies in terms of amount: <img file="imgf000055_0003.tif" frnum="0001" he="13" id="imgf000055_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0050" wi="85"/> It should be mentioned here that the above equation is also a good approximation for the generally considered case of axes of rotation oriented essentially in the same direction. This results in the torque equation (18) for <img file="imgf000055_0002.tif" frnum="0001" he="17" id="imgf000055_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0051" wi="105"/> Here, F<sub>1</sub>, F<sub>2</sub>denote the amounts of the torques produced by the drive devices 1A<sub>1</sub>respectively. 1A<sub>2</sub>generated thrust vectors F<sub>1</sub>, 2001; F<sub>2</sub>, 2002; l<sub>1</sub>, 231 the distance of the thrust vector F<sub>1</sub>, 2001 from the center of mass S, 250 of the aircraft, determined with respect to the first direction (whereby this distance l<sub>1</sub>can be identified with the distance with respect to the first direction between the center of mass S, 250 of the aircraft and the geometric center along the axis of rotation 5A of the propulsion device 1A<sub>1</sub>; in other words: l<sub>1</sub>is the distance with respect to the first direction from the center of mass S, 250 of the aircraft to half the wingspan of the propulsion device 1A<sub>1</sub>); l<sub>2</sub>, 232 the distance of the thrust vector F<sub>2</sub>, 2002 from the center of mass S, 250 of the aircraft, determined with respect to the first direction (whereby this distance l<sub>2</sub>can be identified with the distance with respect to the first direction between the center of mass S, 250 of the aircraft and the geometric center of the propulsion device 1A<sub>2</sub>along the axis of rotation 5A; in other words: l<sub>2</sub>is the distance with respect to the first direction from the center of mass S, 250 of the aircraft to half the span of the propulsion device 1A<sub>2</sub>); F<sub>34</sub>the amount of the thrust vector F<sub>34</sub>= F<sub>3</sub>+ F<sub>4</sub>, 2034 generated by both propulsion devices 1B<sub>3</sub>and 1B<sub>4</sub>; l<sub>34</sub>, 234 the distance, determined with respect to the first direction, between the center of mass S, 250 of the aircraft, on the one hand, and the thrust vector F<sub>34</sub>, 2034 or the axes of rotation of the drive devices 1B<sub>3</sub>and 1B<sub>4</sub>or the straight line that runs through the drive devices 1B<sub>3</sub>and 1B<sub>4</sub>, on the other hand (where it is assumed here that the drive devices 1B<sub>3</sub>and 1B<sub>4</sub>lie on a straight line that runs - at least approximately - parallel to the second direction); a<sub>34</sub>the proportionality factor assigned to the drive devices 1B<sub>3</sub>and 1B<sub>4</sub>.
This equation can now be converted into a ratio of the amounts F<sub>1</sub>and F<sub>2</sub>of the two thrust vectors F<sub>1</sub>, 2001 and 2002 respectively. F<sub>2</sub>, 2002 can be transformed into: (19) <img file="imgf000056_0001.tif" frnum="0001" he="13" id="imgf000056_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0052" wi="49"/> Equation (19) can serve as a configuration formula for the aircraft. Equation (19) initially contains four freely selectable quantities (from the set of F<sub>1</sub>, F<sub>2</sub>, F<sub>34</sub>, l<sub>1</sub>, l<sub>2</sub>l<sub>34</sub>), but in a stable flight attitude equation (17) must also be taken into account, which is why only three of the four quantities mentioned above can be freely selected. A corresponding configuration formula is also obtained for the case that the momentum theorem is set up in the vertical direction 203 (z-axis) and the angular momentum theorem is set up around the first direction 201 (x-axis).
For this purpose, reference is made to Fig. 7c. Such a consideration is necessary if one wants to use the effect according to the invention, i.e. in particular the positive contribution of the Magnus effect, also during forward flight in the second direction (y-axis). The conditions for a stable hover are then: (<sub>20)</sub>(21) <img file="imgf000057_0001.tif" frnum="0001" he="17" id="imgf000057_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0053" wi="100"/> The notations are as in the case of equations (17) and (18), but with the shifted indices: 1 → 3; 2 → 4; 3 → 1; 4 → 2. A repeat of the individual expressions is therefore omitted. In particular, M<sub>12</sub>, 285, is the total torque generated by the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>. Taking into account the explanations in connection with equations (17) and (18) with regard to equation (6), the torque equation (21) can be written as: <img file="imgf000057_0002.tif" frnum="0001" he="17" id="imgf000057_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0054" wi="110"/> Herein, F<sub>3</sub>, F<sub>4</sub>denote the amounts of torque produced by the drive devices 1B<sub>3</sub>and 1B<sub>4</sub>respectively.
1B<sub>4</sub>generated thrust vectors F<sub>3</sub>, 2003; F<sub>4</sub>, 2004 (cf. Fig. 7a); l<sub>3</sub>, 236, the distance of the thrust vector F<sub>3</sub>from the center of mass S, 250 of the aircraft, determined with respect to the second direction (whereby this distance l<sub>3</sub>can be identified with the distance with respect to the second direction between the center of mass S, 250 of the aircraft and the geometric center of the propulsion device 1B<sub>3</sub>along the axis of rotation 5B; in other words: l<sub>3</sub>is the distance with respect to the second direction from the center of mass S, 250 of the aircraft to half the span of the propulsion device 1B<sub>3</sub>); l<sub>4</sub>, 237, the distance of the thrust vector F<sub>4</sub>from the center of mass S, 250 of the aircraft, determined with respect to the second direction (whereby this distance l<sub>4</sub>can be identified with the distance with respect to the second direction between the center of mass S, 250 of the aircraft and the geometric center of the propulsion device 1B<sub>4</sub>along the axis of rotation 5B; otherwise
expressed as: l<sub>4</sub>is the distance in the second direction from the center of mass S, 250 of the aircraft to half the span of the propulsion device 1B<sub>4</sub>); F<sub>12</sub>the amount of the thrust vector F<sub>12</sub>= F<sub>1</sub>+ F<sub>2</sub>, 2012 generated by both propulsion devices 1A<sub>1</sub>and 1A<sub>2</sub>; l<sub>12</sub>, 239, the distance determined in the second direction between the center of mass S, 250 of the aircraft, on the one hand, and the thrust vector F<sub>12</sub>, 2012, or the axes of rotation of the propulsion devices 1A<sub>1</sub>and 1A<sub>2</sub>respectively. the straight line passing through the drive devices 1A<sub>1</sub>and 1A<sub>2</sub>on the other hand (it is assumed here that the drive devices 1A<sub>1</sub>and 1A<sub>2</sub>lie on a straight line which runs - at least approximately - parallel to the first direction); a<sub>12</sub>the proportionality factor assigned to the drive devices 1A<sub>1</sub>and 1A<sub>2</sub>.
This equation can now be converted into a ratio of the amounts F<sub>3</sub>and F<sub>4</sub>of the two thrust vectors F<sub>3</sub>respectively. F<sub>4</sub>can be converted: (22) <img file="imgf000058_0001.tif" frnum="0001" he="17" id="imgf000058_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0055" wi="58"/> Due to the topology of the star-shaped arrangement of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>, 1B<sub>3</sub>and 1B<sub>4</sub>, it is expedient if a pair of drive devices 1A<sub>1</sub>, 1A<sub>2</sub>or 1B<sub>3</sub>, 1B<sub>4</sub>produces half of the required thrust. This results in the boundary condition F<sub>12</sub>= F<sub>34</sub>. (23) It should be noted that this does not necessarily imply that all thrust vectors F<sub>1</sub>, F<sub>2</sub>, F<sub>3</sub>, and F<sub>4</sub> must be equal; it is sufficient if the sum of the thrust vectors of two opposing propulsion devices is equal.
However, all thrust vectors F<sub>1</sub>, F<sub>2</sub>, F<sub>3</sub>, and F<sub>4</sub> can be different when considered individually. A further useful boundary condition arises if it is required that the propulsion devices are preferably mounted centrally on the aircraft fuselage 220. That is, the following applies (24a) (24b) <img file="imgf000058_0002.tif" frnum="0001" he="23" id="imgf000058_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0056" wi="67"/> where for the distance l, 230 of the thrust vectors or the geometric centers of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub> was used: l = l<sub>3</sub>+ l<sub>4</sub>, and for the distance l<sup>'</sup>, 235, the thrust vectors or the geometric centers of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>: l<sup>'</sup>= l<sub>1</sub>+ l<sub>2</sub>.
It is advisable to set l<sup>'</sup>= l. These boundary conditions (23) and (24a) lead to the following configuration formula: (25a) <img file="imgf000059_0001.tif" frnum="0001" he="12" id="imgf000059_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0057" wi="69"/> the boundary conditions (23) and (24b) to: (25b) <img file="imgf000059_0002.tif" frnum="0001" he="12" id="imgf000059_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0058" wi="70"/> It can be conveniently assumed that the proportionality factors are equal, a<sub>12</sub>= a<sub>34</sub>≡ a. Next, the number of freely definable thrust vectors is to be determined using equations (17), (18), (20) and (21).
Assuming that the positions of the drive devices are fixed, the equations in question have the following unknowns: F<sub>1</sub>, F<sub>2</sub>, F<sub>3</sub>, F<sub>4</sub>, l<sub>12</sub>and l<sub>34</sub>. Furthermore, it should be noted that equations (17) and (20) impose the identical constraint. So you have three equations for six unknowns. The center of mass is to be determined using l<sub>12</sub>and l<sub>34</sub>; thus, equations (17), (18), (20) and (21) define another thrust vector; three of the four thrust vectors F<sub>1</sub>, F<sub>2</sub>, F<sub>3</sub>can thus be specified arbitrarily. If further boundary conditions are taken into account, the number of freely definable thrust vectors is reduced accordingly. There are several ways to satisfy equations (17), (20), (25a), (25b). (i) In a first case, it can be required that the aircraft is designed symmetrically.
This means that the center of mass S, 250 is located exactly in the middle between the (centers of mass of the) drive devices 1A<sub>1</sub>, 1A<sub>2</sub>and/or 1B<sub>3</sub>, 1B<sub>4</sub>. In this case l<sub>1</sub>= l<sub>2</sub>and/or l<sub>3</sub>= l<sub>4</sub>. From equations (25a), (25b) it then follows that the propulsion device 1A<sub>1</sub>, 1B<sub>3</sub> arranged in the positive first direction or positive second direction <img file="imgf000060_0003.tif" frnum="0001" he="12" id="imgf000060_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0060" wi="15"/> must generate more thrust than the propulsion device 1A<sub>2</sub>, 1B<sub>4</sub> arranged in the negative first direction or negative first direction, therefore F<sub>1</sub>> F<sub>2</sub>and/or F<sub>3</sub>> F<sub>4</sub>.
Thus, the drive devices arranged in the positive direction must be designed larger than the drive devices arranged in the negative direction. In other words, the propulsion systems located at the front in the forward flight direction must be larger than the propulsion systems located at the rear. In this case, the centre of mass S, 250 will therefore tend to move in the positive first and/or second direction, with the result that l<sub>1</sub>< l<sub>2</sub>and/or l<sub>3</sub>< l<sub>4</sub>, and the difference between the required thrust vectors F<sub>1</sub>and F<sub>2</sub>respectively. F<sub>3</sub>and F<sub>4</sub>of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>respectively. 1B<sub>3</sub>, 1B<sub>4</sub>continues to increase. (ii) In a second case, the two drive devices 1A<sub>1</sub>, 1A<sub>2</sub>are particularly preferably designed to be structurally identical and/or the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub>are particularly preferably designed to be structurally identical.
This means that they are identical in construction and, for example, have the same size, the same wingspan, the same number of rotor blades and/or generate similar or identical (maximum) thrust forces/thrust vectors. In this case, F<sub>1</sub>= F<sub>2</sub>(or F<sub>1</sub>≈ F<sub>2</sub>) and/or F<sub>3</sub>= F<sub>4</sub>(or F<sub>3</sub>≈ F<sub>4</sub>). From equations (25a) and (25b) then follow (26a) (26b) <img file="imgf000060_0001.tif" frnum="0001" he="21" id="imgf000060_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0061" wi="45"/> It can be seen that the center of mass S, 250 of the aircraft is shifted along the first direction 201 and/or second direction from the (geometric) center l/2 between the respective opposite drive devices 1A<sub>1</sub>, 1A<sub>2</sub>or 1B<sub>3</sub>, 1B<sub>4</sub>in the direction of the rear drive devices 1A<sub>2</sub>, 1B<sub>4</sub>with respect to the forward flight direction, namely according to equations (24a) or (24b) by (27a) (27b) <img file="imgf000060_0002.tif" frnum="0001" he="28" id="imgf000060_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0062" wi="64"/> If an aircraft with structurally identical, equally sized drive devices 1A<sub>1</sub>, 1A<sub>2</sub>and/or 1B<sub>3</sub>, 1B<sub>4</sub>and thus approximately equal thrust forces / thrust vectors per pair of propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>respectively.
1B<sub>3</sub>, 1B<sub>4</sub>configured, the center of mass S, 250 can therefore be optimally positioned so that the torques generated by the drive devices are compensated purely by the position of the center of mass S, 250. The said optimal position is determined by equations (27a) and/or (27b). Here and in the following, it must be noted that for the considerations concerning the co-rotating drive devices 1B<sub>3</sub>, 1B<sub>4</sub>, only the position of the center of mass in the first direction 201 plays a role. The storage or Positioning of the center of mass with respect to the second direction and vertical direction 203 is not relevant here and is at the discretion of the expert. Accordingly, for the considerations concerning the co-rotating drive devices 1A<sub>1</sub>, 1A<sub>2</sub>, only the position of the center of mass in the second direction plays a role.
The storage or Positioning of the center of mass with respect to the first direction 201 and vertical direction 203 is not relevant in this case. However, if the aircraft is to exploit the positive effect of the Magnus effect both when moving forward in the first direction and when moving forward in the second direction, the optimal position of the center of mass is determined by both equations (27a) and (27b), so that only its positioning with respect to the vertical direction 203 remains freely selectable. (iii) According to the invention, it is also possible for aspects of the first case design (i) and the second case design (ii) to be combined with one another. This means that the center of mass S, 250 of the aircraft can be displaced from the geometric center of the aircraft fuselage 220 in such a way that the conditions (17), (20), (25a), (25b) for a stable hovering flight are met at certain predetermined, even different, thrust vectors/thrust forces of individual propulsion devices.
For practical applications, it is not always possible to place the masses in an aircraft in such a way that the total center of mass S, 250 can be positioned exactly at the optimal position described in (i), (ii) or (iii) (for case (i) l<sub>1</sub>= l<sub>2</sub>and/or l<sub>3</sub>= l<sub>4</sub>; for case (ii) cf. equations (26a), (26b), (27a), (27b)). Therefore, an area is defined below in which the center of mass S, 250 can lie, so that it is still possible to support the torque compensation with the thrust forces / thrust vectors F<sub>1</sub>, 2001, F<sub>2</sub>, 2002 of the pairs of drive devices 1A<sub>1</sub>, 1A<sub>2</sub>or the torque compensation with the thrust forces / thrust vectors F<sub>3</sub>, 2003, F<sub>4</sub>, 2004 of the pairs of drive devices 1B<sub>3</sub>, 1B<sub>4</sub>.
For this purpose, it is first assumed that one of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>, 1B<sub>3</sub>, 1B<sub>4</sub> has a maximum permissible (i.d. R. predetermined) thrust / maximum permissible thrust vector of F<sub>i,max</sub>. It is assumed that F<sub>i,max</sub>is greater than or equal to the thrust forces F<sub>i,opt</sub>corresponding to the optimal configuration (as already described in more detail in connection with the first aspect of the invention). Taking into account the momentum theorem according to equation (17), the first result is <img file="imgf000062_0001.tif" frnum="0001" he="8" id="imgf000062_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0063" wi="118"/> and thus a maximum permissible thrust vector ratio of <img file="imgf000062_0002.tif" frnum="0001" he="11" id="imgf000062_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0064" wi="42"/> For the case <img file="imgf000062_0003.tif" frnum="0001" he="10" id="imgf000062_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0065" wi="119"/>, the minimum permissible thrust vector ratio is <img file="imgf000062_0004.tif" frnum="0001" he="18" id="imgf000062_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0066" wi="44"/> Using the boundary condition of equation (23), F<sub>12</sub>= F<sub>34</sub>, in equations (17) and (20), the result is <img file="imgf000062_0005.tif" frnum="0001" he="13" id="imgf000062_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0067" wi="54"/> These thrust vector ratios F<sub>1</sub>/F<sub>2</sub> are also described by equation (25a); Using the latter, the maximum permissible distance in the first direction 201 of the centre of mass S, 250 from the geometric centre of the front propulsion device 1A<sub>1</sub> in forward flight, to <img file="imgf000062_0006.tif" frnum="0001" he="19" id="imgf000062_0006" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0068" wi="94"/> and the minimum permissible distance in the first direction 201 of the centre of mass S, 250 from the geometric centre of the front propulsion device 1A<sub>1</sub> in forward flight, to can be calculated.
<img file="imgf000063_0001.tif" frnum="0001" he="16" id="imgf000063_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0069" wi="97"/> If the centre of mass S, 250 lies outside the range <img file="imgf000063_0002.tif" frnum="0001" he="12" id="imgf000063_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0070" wi="46"/> (28) with respect to the first direction 201, it is no longer possible to compensate for the deviation of the centre of mass S, 250 from the optimum position according to equation (26a) by the thrust forces F<sub>1</sub>, 2001 or F<sub>2</sub>, 2002 of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>. By means of equation (24a), the permissible range (28) in the first direction can also be specified with respect to the axes of rotation of the drive devices 1B<sub>3</sub>, 1B<sub>4</sub>or the straight line passing through the drive devices 1B<sub>3</sub>, 1B<sub>4</sub>.
Then the area is specified using the distance l<sub>34</sub>and corresponding limits l<sub>34,min</sub>and l<sub>34,max</sub>. Analogously, one obtains for the permissible range of the center of mass S, 250 with respect to the second direction (here: y-direction) (29) where <img file="imgf000063_0003.tif" frnum="0001" he="74" id="imgf000063_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0071" wi="102"/> Using equation (24b), the permissible range (29) in the second direction can also be specified with respect to the axes of rotation of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>or the straight line that runs through the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>. Then the Specify the area using the distance l<sub>12</sub>and corresponding limits l<sub>12,min</sub>and l<sub>12,max</sub>.
Figure 7d serves to illustrate the region described above in which the center of mass S, 250 of the aircraft can expediently be located for implementing the invention according to the second aspect. Fig. 7d shows schematically an aircraft with propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>and 1B<sub>3</sub>, 1B<sub>4</sub>, which corresponds to that described in connection with Figures 7a and 7b. It is further assumed that the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>1B<sub>3</sub>, 1B<sub>4</sub> are structurally identical (see Case (ii) above), here in particular: a<sub>1</sub>= a<sub>2</sub>= a<sub>3</sub>= a<sub>4</sub>= a<sub>12</sub>= a<sub>34</sub>≡ a. First, it is further assumed that the torque compensation is to be realized purely via the position of the center of mass S, 250, whereby <img file="imgf000064_0002.tif" frnum="0001" he="10" id="imgf000064_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0072" wi="102"/> applies.
For the embodiment of the aircraft considered here, a total weight force of F<sub>s</sub>= 1000 ^ generated by a corresponding total mass is assumed; the characteristic number / the proportionality factor is typically a = 0.2 m; the distance of the drive devices in the first direction (in Figures 7a, 7b: x-direction) is defined as l = l<sub>1</sub>+ l<sub>2</sub>= 2 m. Based on these specifications, equations (25a) and (26a) result in an optimal center of mass position of <img file="imgf000064_0001.tif" frnum="0001" he="39" id="imgf000064_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0073" wi="120"/>. If it is not possible to set the total center of mass S, 250 of the aircraft to the position l<sub>1,opt</sub>= 1.1 m, a range is now defined in which the position of the
center of mass S, 250, so that the torque compensation can be compensated by the thrust forces / thrust vectors of the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>, 1B<sub>3</sub>, 1B<sub>4</sub>. For this purpose, the maximum permissible thrust that can be generated by each of the first-direction drive devices 1A<sub>1</sub>, 1A<sub>2</sub>, which is suitably controlled by the thrust vector control, is defined as <img file="imgf000065_0003.tif" frnum="0001" he="9" id="imgf000065_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0074" wi="51"/>. By this specification and using the boundary condition according to equation (23), , the maximum and minimum permissible thrust vector ratio <img file="imgf000065_0004.tif" frnum="0001" he="10" id="imgf000065_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0075" wi="33"/> <img file="imgf000065_0001.tif" frnum="0001" he="26" id="imgf000065_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0076" wi="48"/> and the range for the position of the center of mass according to equation (28) <img file="imgf000065_0002.tif" frnum="0001" he="33" id="imgf000065_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0077" wi="106"/> can be calculated.
This means that in this example the centre of mass with respect to the first direction is conveniently located 1.05 to 1.15 m from the geometric centre of the front of the two propulsion devices 1A<sub>1</sub> with respect to the forward flight direction. Using equation (27a), this result can also be expressed as follows: the center of mass is conveniently located 0.05 to 0.15 m from the axis of rotation of the drive devices 1B<sub>3</sub>, 1B<sub>4</sub> or the straight line passing through the two drive devices 1B<sub>3</sub>, 1B<sub>4</sub> with respect to the first direction. Assuming that the aircraft is designed symmetrically, the same values are obtained for the range permitted for l<sub>3</sub>. Taking both conditions into account, the centre of mass S, 250 is conveniently positioned with respect to the plane defined by the propulsion devices and the aircraft fuselage within a quadratic region determined by the limits given.
Positioning in the vertical direction is not restricted. Finally, it is stated that the second aspect of the invention is not limited to aircraft with four propulsion devices. It is also possible, for example, for more than two drive devices to be arranged along one direction or for some drive devices to be arranged on straight lines parallel to one another. The equations (17), (18), (20), (21) are now generalized for an aircraft according to the invention with n, n > 2, propulsion devices 1C. Figure 8a shows a section of such an aircraft in plan view; Figure 8b shows a section of the aircraft in side view. We assume that the mathematical-physical description of the aircraft is done in a Cartesian coordinate system with x-, y- and z-axes. The n propulsion devices 1C and the aircraft fuselage 220 are located in the xy plane, i.e. in the plane with z = 0. The propulsion devices 1C are arranged around the aircraft fuselage 220 (star-shaped) in the plane z = 0.
The origin O of the coordinate system lies in the geometric center of the aircraft. Then let r<sub>i</sub>, i ∈ {1, … ,n} be the position vectors to the i-th thrust vector of the corresponding propulsion devices 1C. Let s be the position vector to the center of mass S, 250 of the aircraft. The vector of the aircraft's weight force is F<sub>s</sub>= (0,0,F<sub>s</sub>). In the case of stable hovering flight of interest here, the thrust vectors <img file="imgf000066_0005.tif" frnum="0001" he="7" id="imgf000066_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0078" wi="28"/> , , generated by the propulsion devices are: <img file="imgf000066_0003.tif" frnum="0001" he="9" id="imgf000066_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0079" wi="71"/> Finally, in hovering flight, the propulsion devices rotate with the angular velocity <img file="imgf000066_0004.tif" frnum="0001" he="7" id="imgf000066_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0080" wi="28"/> where these are vectors that lie in the xy plane.
The torque that must be applied by the aircraft, already described in detail in the introduction, is then, taking into account the relationship M<sub>i</sub>= a<sub>i</sub>* F<sub>i</sub>: (<sub>30)</sub> <img file="imgf000066_0002.tif" frnum="0001" he="12" id="imgf000066_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0081" wi="77"/> The equilibrium conditions of equations (17), (18), (20), (21) are then: (<sub>31)</sub>(32) <img file="imgf000066_0001.tif" frnum="0001" he="19" id="imgf000066_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0082" wi="84"/> “×” denotes the cross product. From the angular momentum theorem, the position vector s of the center of gravity S, 250 can be determined as follows: With <img file="imgf000067_0001.tif" frnum="0001" he="19" id="imgf000067_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0083" wi="95"/> using the Grassmann identity: <img file="imgf000067_0002.tif" frnum="0001" he="8" id="imgf000067_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0084" wi="122"/> and taking into account that F<sub>i</sub> is always normal to (r<sub>i</sub>- s), whereby their scalar product is zero: <img file="imgf000067_0003.tif" frnum="0001" he="9" id="imgf000067_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0085" wi="94"/> one first obtains <img file="imgf000067_0004.tif" frnum="0001" he="23" id="imgf000067_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0086" wi="122"/> and finally the position vector s of the center of gravity S, 250: (<sub>33)</sub> <img file="imgf000067_0005.tif" frnum="0001" he="13" id="imgf000067_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0087" wi="56"/> Equation (32) provides two conditions, namely one for the x-components of the torques, and another for their y-components.
Equations (31) and (32) (taking into account the context of equation (30)) thus provide three conditions for the magnitudes F<sub>i</sub>of the n thrust vectors and the two coordinates of the center of gravity. This leaves n + 2 – 3 = n – 1 of the n thrust vectors that can be specified. A suitable range with respect to the plane in which the propulsion devices 1C and the aircraft fuselage 220 are located can thus also be determined in the more general case considered here by varying the thrust vectors of one or more of the n propulsion devices 1C and requiring that the center of mass S, 250 must be positioned in such a way (cf. Equation (33)) that the torque compensation according to equation (32) can be compensated by the thrust forces / thrust vectors of the drive devices. For this purpose, it may be appropriate to drive one or more of the drive devices with the maximum permissible thrust. Since the optimal position of the centre of mass in the configuration under consideration is determined by the intersection of two straight lines, it is convenient to consider the first direction and/or the second direction along which the propulsion devices rotate substantially in the same direction as the directions perpendicular to two given forward flight directions.
In this case, the The center of mass is thus preferably (i) displaced from the geometric center with respect to a direction perpendicular to the first direction and lying in the plane defined by the propulsion devices and the aircraft fuselage, and/or (ii) displaced from the geometric center with respect to a direction perpendicular to the second direction and lying in the plane defined by the propulsion devices and the aircraft fuselage. Figure 9a shows an embodiment according to the second aspect of the invention, in which three drive devices 1C<sub>1</sub>, 1C<sub>2</sub>, 1C<sub>3</sub> are arranged around the aircraft fuselage 220 of the aircraft in such a way that they form the corners of an equilateral triangle. It is shown that the drive devices 1C<sub>1</sub>and 1C<sub>2</sub>are arranged on a straight g<sub>1</sub>; g<sub>1</sub>defines the first direction according to the invention.
In the embodiment shown, the drive device 1C<sub>3</sub>is arranged on a straight line g<sub>2</sub>, which is perpendicular to the straight line g<sub>1</sub> and runs through the geometric center G of the aircraft, in this case through the geometric center G of the equilateral triangle. The straight line g<sub>2</sub>defines the second direction according to the invention. The rotary axes 5C<sub>1</sub>, 5C<sub>2</sub>, 5C<sub>3</sub>of the drive devices 1C<sub>1</sub>, 1C<sub>2</sub>respectively. 1C<sub>3</sub>point towards (or away from) the geometric center G. In the embodiment shown, only the axis of rotation 5C<sub>3</sub>is aligned exactly parallel to the second direction defined by g<sub>2</sub>. The rotation axes 5C<sub>1</sub>, 5C<sub>2</sub>are not exactly parallel to the first direction defined by g<sub>1</sub>.
As can be seen from simple geometric considerations, the axis of rotation 5C<sub>1</sub>encloses an angle Į<sub>1</sub>= 30° with the straight line g<sub>1</sub>(first direction); likewise, the axis of rotation 5C<sub>2</sub>encloses an angle Į<sub>2</sub>= 30° with the straight line g<sub>1</sub>(first direction). Such angles fall under the inventive concept of axes of rotation aligned substantially in the first direction. Preferably, the angles can also be chosen smaller. If the drive devices 1C<sub>1</sub>, 1C<sub>2</sub> rotate about these axes of rotation essentially in the same direction of rotation as defined above, the effect according to the invention also occurs here when the aircraft moves in particular along the second direction defined by g<sub>2</sub>. Rotate the drive devices 1C<sub>1</sub>, 1C<sub>3</sub>about the corresponding rotation axes 5C<sub>1</sub>resp.
5C<sub>3</sub>essentially in the same direction of rotation, the advantage according to the invention has a positive effect, particularly in the case of a movement along the angle bisector 1C<sub>1</sub>-G-1C<sub>3</sub>. Figure 9b shows an aircraft according to the second aspect of the invention, in which seven propulsion devices 1C<sub>1</sub>, ..., 1C<sub>7</sub> are arranged in a plane around the aircraft fuselage 220. The drive devices 1C<sub>1</sub>, …, 1C<sub>7</sub> are arranged such that they form the corners of a regular heptagon. Each of the drive devices is mounted so as to be rotatable about an associated axis of rotation 5C<sub>1</sub>, …, 5C<sub>7</sub>. In the embodiment shown, the rotation axes 5C<sub>1</sub>, …, 5C<sub>7</sub>point to the geometric center G of the aircraft or the heptagon. This embodiment is intended to describe the general case in which (an odd number) n = 2j + 1, j > 1, propulsion devices 1C<sub>1</sub>, ..., 1C<sub>2j + 1</sub> are arranged around the aircraft fuselage 220 such that they form the corners of a regular (2j + 1)-corner.
The corresponding rotation axes 5C<sub>1</sub>, …, 5C<sub>2j + 1</sub>should point towards (or away from) the geometric center G. In this case, it is expedient to consider a first straight line g<sub>1</sub> which passes through two drive devices 1C<sub>1</sub>and 1C<sub>(n + 1)/2</sub>; this straight line g<sub>1</sub> defines the first direction according to the invention. Furthermore, it is expedient to consider a second straight line g<sub>2</sub>which passes through two drive devices 1C<sub>k</sub>and 1C<sub>k + (n – 1)/2</sub>, <img file="imgf000069_0002.tif" frnum="0001" he="9" id="imgf000069_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0088" wi="56"/>rounding its argument to the nearest whole number); this straight line g<sub>2</sub>defines the second direction according to the invention.
Using simple geometric considerations, it can be seen that each of the rotation axes 5C<sub>1</sub>and 5C(n + 1)/2 encloses an angle a1 = a(n+1)/2 =90°/n with the straight line g1 (i.e. the first direction); the same applies to the angles between the rotation axes 5Ck and 5C<sub>k + (n – 1)/2</sub>and the straight line g<sub>2</sub>: a<sub>k</sub>= a<sub>k+(n-1)/2</sub>= 90°/n. In the case of the heptagon shown, a<sub>1</sub>= a<sub>3</sub>= a<sub>4</sub>= a<sub>6</sub>= 90°/7 ^ 12.86°. For a regular (2j + 1)-gon, it is therefore advantageous if the axes of rotation of the drive devices, which lie on the straight lines g<sub>1</sub>and g<sub>2</sub>defining the first and second directions, enclose an angle between 0° and 90°/n with the associated straight lines g<sub>1</sub>and g<sub>2</sub>. The angle İ between g<sub>1</sub>and g<sub>2</sub>is given by as is easily deduced from <img file="imgf000069_0001.tif" frnum="0001" he="12" id="imgf000069_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0089" wi="31"/> using geometric relations.
Figure 9c shows an aircraft according to the second aspect of the invention, in which six propulsion devices 1C<sub>1</sub>, ..., 1C<sub>6</sub> are arranged in a plane around the aircraft fuselage 220. The drive devices 1C<sub>1</sub>, …, 1C<sub>6</sub> are arranged such that they form the corners of a regular hexagon. Each of the drive devices is mounted so as to be rotatable about an associated axis of rotation 5C<sub>1</sub>, …, 5C<sub>6</sub>. In the embodiment shown the rotation axes 5C<sub>1</sub>, …, 5C<sub>6</sub>point to the geometric center G of the aircraft or the hexagon. This embodiment is intended to describe the general case in which (an even number) n = 2j, j > 1, drive devices 1C<sub>1</sub>, ..., 1C<sub>2j</sub> are arranged around the aircraft fuselage 220 such that they form the corners of a regular 2j-gon.
The aircraft fuselage 220 is located between two opposite propulsion devices of the regular 2j-corner. The corresponding rotation axes 5C<sub>1</sub>, …, 5C<sub>2j</sub>should point towards (or away from) the geometric center G. In this case, it is expedient to consider a first straight line g<sub>1</sub> which passes through two drive devices 1C<sub>1</sub>and 1C<sub>n /2 + 1</sub>; this straight line g<sub>1</sub> defines the first direction according to the invention. Furthermore, it is expedient to consider a second straight line g<sub>2</sub>which runs through two drive devices 1C<sub>k</sub>and 1C<sub>k +n /2</sub>, <img file="imgf000070_0001.tif" frnum="0001" he="9" id="imgf000070_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0090" wi="31"/> this straight line g<sub>2</sub>defines the second direction according to the invention.
In the embodiment shown, the axes of rotation of the drive devices arranged on the straight lines g<sub>1</sub>and g<sub>2</sub>are aligned (mathematically exactly) parallel in the first and second directions, respectively. Particularly preferably, the first and second directions are substantially perpendicular, especially perpendicular, to each other; this is always possible when the drive devices form the corners of a 4j-gon. The angle İ between g<sub>1</sub>and g<sub>2</sub>(i.e. first and second direction) is given for the 2j-corner described above by <img file="imgf000070_0002.tif" frnum="0001" he="9" id="imgf000070_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0091" wi="50"/> as can easily be seen by using geometric relations. From the above embodiments it can be seen that for the arrangement of any (even or odd) number of drive devices at the corners of a regular n-gon it is sufficient if the axes of rotation of the drive devices, which lie on the straight lines g<sub>1</sub>and g<sub>2</sub> defining the first and second directions, enclose an angle between 0° and 30° (for n > 2), particularly preferably between 0° and 18° (for n > 3) with the associated straight lines g<sub>1</sub>and g<sub>2</sub>; furthermore it is expedient if the straight lines g<sub>1</sub>and g<sub>2</sub>(and thus the first and second directions) are selected such that the angle between them is greater than or equal to 60°, especially in the range between 60° and 90°.
Appendix (Derivation of the relationship between thrust and power) The derivation of thrust and power is based on the jet theory, whereby a drive device / rotor is considered as an actuator disk without information about the number and shape of the rotor blades. The flow is simplified as one-dimensional, quasi-stationary, incompressible and frictionless, which results in the corresponding conservation laws for mass, momentum and energy. In the following, all sizes in the actuator disk plane are marked with the additional index a, all sizes far above the actuator disk plane (inflow plane) with the additional index 0 and all sizes far below the actuator disk plane (outflow plane) with the additional index λ. Law of conservation of mass: Based on the assumptions regarding the flow, the mass flow follows from the law of conservation of mass: <img file="imgf000071_0001.tif" frnum="0001" he="46" id="imgf000071_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0092" wi="126"/> Law of conservation of momentum: Based on the assumptions regarding the flow, the thrust follows from the law of conservation of momentum: <img file="imgf000071_0002.tif" frnum="0001" he="25" id="imgf000071_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0093" wi="109"/> Since the rotor does not influence the inflow plane, v<sub>i0</sub>= 0, from which <img file="imgf000071_0003.tif" frnum="0001" he="9" id="imgf000071_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0094" wi="34"/> follows.
By inserting the mass flow in the actuator disk plane we get: <img file="imgf000072_0001.tif" frnum="0001" he="16" id="imgf000072_0001" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0095" wi="69"/> Law of conservation of energy: Based on the assumptions regarding the flow and v<sub>i0</sub>= 0, the law of conservation of energy gives the power or work done per unit of time for the actuator disk plane: <img file="imgf000072_0002.tif" frnum="0001" he="17" id="imgf000072_0002" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0096" wi="163"/> By inserting the mass flow in the actuator disk plane we get: <img file="imgf000072_0003.tif" frnum="0001" he="14" id="imgf000072_0003" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0097" wi="98"/> Using the thrust force we get the power as: <img file="imgf000072_0004.tif" frnum="0001" he="16" id="imgf000072_0004" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0098" wi="89"/> These equations give us the relationship <img file="imgf000072_0005.tif" frnum="0001" he="13" id="imgf000072_0005" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0099" wi="45"/> directly, whereby the power can be expressed as <img file="imgf000072_0006.tif" frnum="0001" he="15" id="imgf000072_0006" img-content="drawing" img-format="tif" inline="no" orientation="portrait" pgnum="0100" wi="102"/>.
[0002]
List of reference numerals 100 Aircraft according to the first aspect of the invention 120 Aircraft body 1F Drive devices arranged in the front area 1R Drive devices arranged in the rear area 101 Longitudinal direction of the aircraft 100 102 Transverse direction of the aircraft 100 103 Vertical direction of the aircraft 100 121 Bow / nose of the aircraft 100 122 Rear of the aircraft 100 1 Drive device 2 Rotor blades of a drive device 3 Pitch mechanism 31 Coupling device 32 Coupling point 33 Bearing device 4 Offset device 11 Disk of the drive device 1 5 Axis of rotation of a drive device 51 Direction of rotation of a drive device 52 Radius of the drive device 61 Connecting element 7, 71 Force on a drive device / thrust vector 72 Contribution of the Magnus effect to the thrust vector 8 Torque on a drive device 9 Air flow 110 Arrow indicating the direction of movement of the aircraft 150 Center of mass of the aircraft 100 701 Total thrust vector generated by the propulsion devices 1F 702 Total thrust vector generated by the propulsion devices 1R 81 Total torque generated by the propulsion devices 1F 82 Total torque generated by the propulsion devices 1R
131 Distance in the longitudinal direction between the center of mass 150 and the propulsion devices 1F 132 Distance in the longitudinal direction between the center of mass 150 and the propulsion devices 1R 160 Weight of the aircraft g<sub>i</sub>i-th straight line along which the propulsion devices are arranged n<sub>i</sub>number of propulsion devices arranged along the straight line g<sub>i</sub>N total number of straight lines K total number of propulsion devices F<sub>ij</sub>thrust vector generated by the j-th propulsion device arranged on the straight line gi F<sub>i</sub>thrust vector generated by all the propulsion devices arranged on a straight line gi M<sub>i</sub>torque generated by all the propulsion devices arranged on a straight line gi x<sub>i</sub>longitudinal coordinate of the straight line g<sub>i</sub>X<sub>S</sub>longitudinal coordinate of the center of mass 150 200 Aircraft according to the second aspect of the invention 220 Aircraft fuselage 1A, 1B, 1C, 1A<sub>1</sub>, 1A<sub>2</sub>, 1B<sub>3</sub>, 1B<sub>4</sub>, 1C<sub>i</sub>Drive devices of the aircraft 200 221, 222 Arms for coupling the drive devices 1A, 1B to the aircraft fuselage 220 201 first direction 202 second direction 203 vertical direction 5A rotation axes of the drive devices 1A 5B rotation axes of the drive devices 1B 5C<sub>i</sub>rotation axis of the drive device 1C<sub>i</sub>α<sub>i</sub>angle between rotation axis 5C<sub>i</sub>and first or second direction ε angle between first and second direction 250 Center of mass of the aircraft 200 G geometric center O origin of the coordinate system 2001, 2002 thrust vectors generated by the propulsion devices 1A<sub>1</sub>or 1A<sub>2</sub>
2003, 2004 Thrust vectors generated by the propulsion devices 1B<sub>3</sub>resp. 1B<sub>4</sub> are generated 2012 total thrust vector generated by the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>2034 total thrust vector generated by the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub> 230 distance of the thrust vectors / geometric centers of the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub> 231 distance of the thrust vector 2001 from the center of mass 250 of the aircraft 232 distance of the thrust vector 2002 from the center of mass 250 of the aircraft 234 distance between the center of mass 250 and the thrust vector F<sub>34</sub>, 2034 / the axes of rotation of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>/ the straight line passing through the Propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>runs 235 Distance of the thrust vectors / geometric centers of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>236 Distance of the thrust vector 2003 from the center of mass 250 of the aircraft 237 Distance of the thrust vector 2004 from the center of mass 250 of the aircraft 239 Distance between the center of mass 250 and the thrust vector F<sub>12</sub>, 2012 / the axes of rotation of the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>/ the straight line that runs through the propulsion devices 1A<sub>1</sub>, 1A<sub>2</sub>runs 251 Direction of rotation of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>280 total of the propulsion devices 1B<sub>3</sub>, 1B<sub>4</sub>generated torque 285 total torque generated by the drive devices 1A<sub>1</sub>, 1A<sub>2</sub>