aircraft

Abstract

The present invention relates to an aircraft (100) including an aircraft body (120) defining a longitudinal direction, a vertical direction and a transverse direction, and at least two propulsion devices (1) rotatable about respective associated rotation axes (5) and generating respective associated thrust vectors, a first number of the propulsion devices being arranged along a first straight line parallel to the transverse direction, a second number of the propulsion devices being arranged along a second straight line parallel to the transverse direction, the first straight line being spaced apart from the second straight line, and the aircraft's center of gravity being arranged between the first straight line and the second straight line. The aircraft is adapted to perform hover flight, during which each of the associated rotation axes (5) is substantially oriented in a transverse direction of the aircraft body, and each of the at least two propulsion devices rotates in substantially the same rotational direction about each associated rotation axis. Aircraft configurations with further orientations of the rotation axes are also conceivable.

Description

[Technical Field]

[0001] The present invention relates to an aircraft and to a method for manufacturing and controlling an aircraft, and in particular to an aircraft capable of performing stable hovering flight by means of propulsion devices rotating in the same direction, in particular cyclogyro rotors. [Background technology]

[0002] Aircraft that use cyclogyro rotors as their propulsion device are called cyclogyros. Like helicopters, cyclogyros are also vertical take-off and landing (VTOL) aircraft, meaning they can take off and land vertically without a runway.

[0003] The cyclogyro rotor is based on the principle of thrust generation by rotating wings, hereafter referred to as rotor blades. In contrast to classical rotors, such as those used in helicopter propulsion systems, the rotation axis of the blades of a cyclogyro rotor is aligned parallel to the longitudinal axis of the blade / rotor blade. The direction of thrust of the entire cyclogyro rotor is perpendicular to the rotation axis.

[0004] In both cases, during stationary maneuvers such as hovering or constant forward flight, all rotor blades of a cyclogyro rotor should ideally be oriented in the best possible direction relative to the flow direction at all times to maximize their contribution to the total thrust with the minimum thrust required. The maximum pitch of the rotor blades relative to the flow direction directly affects the amount of thrust generated. As the rotor rotates, the pitch of each rotor blade must be continuously changed during one revolution. Thus, each rotor blade of a cyclogyro rotor undergoes a periodic change in pitch angle. This periodic change in pitch angle is called pitch motion.

[0005] Various pitch mechanisms are known for generating pitch motion. For example, each rotor blade may be connected to an eccentric bearing shaft via one or more connecting rods. The resulting pitch motion of the rotor blades is cyclically repeated with each rotor rotation.

[0006] Various embodiments of propulsion devices for cyclogyros are described, for example, in EP 3548378 A1 and EP 3715249 A1.

[0007] Periodic adjustment of the rotor blades generates a thrust vector perpendicular to the rotor's axis of rotation. An offset device is used to periodically change the rotor blade pitch, so that the thrust vector can be rotated in the entire plane perpendicular to the rotor's axis of rotation (thrust vector control). In addition to the thrust vector, the rotor generates a torque about the axis of rotation that is opposite to the direction of the rotor's rotation. This torque results from the tangential components of the aerodynamic forces, i.e., lift and drag, acting on the rotor blades.

[0008] When air flows outward relative to the rotor, the aerodynamic properties, and therefore the characteristics of the thrust vector produced, change. When the rotor is in forward flight, air is actively blown onto the rotor from the front. The changing properties can be roughly explained by the Magnus effect, which states that a sphere rotating in a flow experiences a lateral force perpendicular to the direction of the flow.

[0009] The direction of the lateral force depends on the direction of rotation of the body, or in this case the cyclogyro rotor.

[0010] However, in aircraft or cyclogyros known from, for example, the paper "Development of a Four-Rotor Cyclocopter" by I. S. Wang et al., Journal of Aircraft, Vol. 45, No. 6, November-December 2008, pages 2151 ff., and the paper "Experimental Optimization of MAV-Scale Cycloidal Rotor Performance" by M. Benedict et al., Journal of the American Helicopter Society, Vol. 56, 2005 (2011), the rotors rotate in opposite directions while the incident flow remains constant. In this case, i.e., when the rotors rotate in opposite directions, the rotor side forces caused by the Magnus effect do not act in the same direction. Therefore, while the same lift is required, total thrust may decrease or power requirements may increase. Therefore, at higher forward speeds and opposite rotational directions, the negative effects of the Magnus effect may no longer be compensated for by the rotors. As a result, the aircraft can no longer fly, and the rotors cannot be used as lift-generating components. Summary of the Invention [Problem to be solved by the invention]

[0011] SUMMARY OF THE INVENTION It is therefore an object of the present invention to provide an aircraft that can assume a stable flight attitude in forward flight even at high speeds. [Means for solving the problem]

[0012] This problem is solved by an aircraft having the features set forth in claim 1, an aircraft having the features set forth in claim 5, a method for manufacturing an aircraft set forth in claims 17 and 18 respectively, and a method for controlling an aircraft set forth in claims 19 and 20 respectively. Advantageous embodiments of the invention are disclosed in the dependent claims 2-4, 6-16 and 21-24.

[0013] According to a first aspect of the present invention, there is provided an aircraft including an aircraft body defining a longitudinal direction, a vertical direction, and a transverse direction, the longitudinal direction corresponding to the tail-to-nose direction of the aircraft, the vertical direction corresponding to the direction of the Earth's gravitational force when the aircraft is resting on the ground, and the transverse direction perpendicular to the longitudinal and vertical directions, and at least two propulsion devices rotatable about respective axes of rotation and generating respective thrust vectors. A first number of propulsion devices are arranged along a first line parallel to the transverse direction, and a second number of propulsion devices are arranged along a second line parallel to the transverse direction. The first line is spaced from the second line, and the center of gravity of the aircraft is located between the first line and the second line with respect to the longitudinal direction. The aircraft is adapted to perform hover flight whereby all forces acting on the aircraft and all torques acting on the aircraft substantially disappear relative to the aircraft's center of gravity, and during hover flight each of the associated rotational axes is oriented substantially transverse to the aircraft body, and each of the at least two propulsion devices rotates in substantially the same rotational sense about its respective associated rotational axis.

[0014] According to the invention, an axis of rotation is oriented substantially in the transverse direction of the aircraft body when the angle included between the axis of rotation and an axis extending transversely and intersecting the axis of rotation is less than 45°, preferably less than 30°, more preferably less than 15°.

[0015] Therefore, for the purposes of the present invention, it is not necessary that all rotation axes be mathematically precisely parallel during hover flight, and in fact it may be advantageous if the angles between the rotation axes and the transverse axes intersecting them are in the range of 5° to 30°, particularly preferably 10° to 20°.

[0016] Furthermore, according to the present invention, the propulsion devices are considered to be rotating in substantially the same direction of rotation if the scalar product of the vector of the angular velocity of a given propulsion device with a fixed, but arbitrary, vector in the transverse direction has the same sign for all propulsion devices. This means that the transverse vector is first fixed to ensure that all or each of the propulsion devices under consideration are rotating in substantially the same direction of rotation. Then, for a first propulsion device, the scalar product of its angular velocity vector with the fixed vector is calculated, then for a second propulsion device, the scalar product of its angular velocity vector with the fixed vector is calculated, and so on. Finally, only the signs (positive or negative) of the scalar products thus calculated are compared. If all signs are the same, the or each of the propulsion devices under consideration are rotating in substantially the same direction of rotation within the meaning of the present invention.

[0017] Thus, for the purposes of the present invention, during hover flight, it is not necessary that all rotational axes be mathematically exactly parallel, nor that all propulsion devices rotate about their rotational axes at the same rotational or angular velocity (in terms of magnitude).

[0018] By designing the aircraft to perform hover flight using propulsion devices that rotate in substantially the same direction, a reduction in the power consumption of the propulsion devices is achieved. Simply put, the Magnus effect generated in accordance with the present invention replaces part of the thrust of the propulsion devices, thus reducing the power requirements during forward flight compared to hover flight. This therefore leaves more residual power for the propulsion devices during forward flight, thereby particularly improving the agility of the aircraft during forward flight.

[0019] The Magnus effect indicates that a sphere rotating in a flow experiences a lateral force perpendicular to the direction of the flow. In the case of a propulsion device according to the present invention rotating in substantially the same direction, this effect can generate an additional thrust vector or force in the vertical direction. Thus, the total lift of the propulsion device increases. The Magnus effect replaces part of the thrust applied by the propulsion device, thus reducing the power requirement during forward flight compared to hover flight. Here, when the rotor is in forward flight, air actively flows against it from ahead. In the configuration according to the present invention, the inflow remains constant, but with a propulsion device rotating in substantially the same direction, the additional lateral force of the Magnus effect acts in substantially the same direction as the thrust of the propulsion device, thereby increasing the total thrust or reducing the power requirement while requiring the same lift. Therefore, particularly in forward flight at higher forward speeds and with substantially the same rotation direction, the positive effect of the Magnus effect can reduce the power and / or rotation speed that the propulsion device may require to maintain the aircraft in a stable flight attitude.

[0020] In a particularly preferred embodiment, the aircraft is further adapted to be positioned such that, during hover flight, when one or more of the propulsion devices generate a particular predetermined thrust vector associated therewith, the center of gravity of the aircraft is positioned such that all forces acting on the aircraft and all torques acting on the aircraft are substantially nullified relative to the center of gravity of the aircraft. Associated with this indication is the constraint that the center of gravity of the aircraft in the longitudinal direction must be within a range determined by the fact that the aircraft is capable of hovering when one or more of the propulsion devices are driven with maximum thrust or maximum thrust vector. In other words, if the center of gravity is within said range, the propulsion devices are capable of generating an appropriate thrust vector to enable the aircraft to perform hover flight.

[0021] During hover flight, the approach speed is generally lower than during forward flight. Since the thrust vector of the propulsion devices for hover flight is specified for the aircraft according to the invention and the position of the center of gravity for hover flight is determined, it is ensured that a stable flight attitude is also possible during forward flight. As mentioned above, the higher the approach speed, the greater the positive effect caused by the Magnus effect according to the invention. Therefore, the configuration of the aircraft during hover flight according to the invention ensures that the aircraft can assume a stable flight attitude, especially during forward flight, since during forward flight the increase in the thrust vector due to the Magnus effect is greater than in hover flight.

[0022] When designing and configuring an aircraft having a propulsion device according to the invention, all forces and torques of the propulsion device must be taken into account. Essentially, the thrust forces or thrust vectors are used to generate the required lift and / or control the aircraft's flight attitude. To this end, the aircraft advantageously includes a thrust vector control system that adjusts the required thrust forces or required thrust vectors during hover flight and / or forward flight.

[0023] Each propulsion device according to the present invention generates a torque in the direction of rotation. This torque about the axis of rotation in the direction of rotation of the propulsion device results, inter alia, from tangential aerodynamic forces caused by air resistance. Therefore, to maintain a constant rotational speed, the propulsion device must generate a (driving) torque that counteracts the torque generated by the tangential aerodynamic forces. However, to enable the propulsion device to generate such a (driving) torque during the flight phase, another torque is required, which the aircraft body must apply (according to the law of action and reaction) to "support" the propulsion device in the air. This latter torque is approximately equal in magnitude (ignoring dissipation effects) to the torque generated by the tangential aerodynamic forces and is directed in the same direction as the torque generated by the tangential aerodynamic forces in order to maintain a constant rotational speed against the aerodynamic forces. Since the torque generated by the aerodynamic forces is opposite to the direction of rotation of the propulsion device, the torque applied by the aircraft body is also opposite to the direction of rotation of the propulsion device. Assuming that the torque due to the aerodynamic forces and the torque due to the propulsion device are substantially equal in magnitude but directed in opposite directions, the net torque remaining due to the rotation of the propulsion device is the torque applied by the aircraft body.

[0024] According to the invention, this torque or these torques are compensated for by taking into account the thrust vectors assigned to each propulsion device and by positioning the center of gravity of the aircraft such that all forces acting on the aircraft and all torques acting on the aircraft substantially disappear with respect to the center of gravity of the aircraft during hovering flight.

[0025] According to the invention, since the propulsion devices rotate in substantially the same direction, the torques of all these propulsion devices generated by the aircraft body described above also act in substantially the same direction, so that the torques are additive and do not cancel each other out.

[0026] To be able to achieve a stable flight attitude in forward flight as well as in hovering flight, balance of all forces and torques acting on the aircraft must be achieved. This calculation is done using the momentum theorem and the relationship between angular momentum and torque.

[0027] The momentum theorem is,

number

number

[0028] The relationship between angular momentum and torque is

number

number

[0029] For requirements of stable flight attitudes (hovering flight, uniform speed in forward flight, etc.), the acceleration vector

number

number

[0030] Thus, force and torque balance can be achieved by selecting the thrust or thrust vectors of the propulsion devices and their corresponding distances from the center of gravity of the aircraft.

[0031] Preferably, the first number of propulsion devices are arranged in a longitudinally forward region of the aircraft, and the second number of propulsion devices are arranged in a longitudinally aft region of the aircraft. Preferably, the aircraft includes three propulsion devices. Particularly preferably, the aircraft includes four propulsion devices, two of which are arranged in a longitudinally forward region of the aircraft, and the other two are arranged in a longitudinally aft region of the aircraft. The overall length of the aircraft is measured longitudinally. To simplify the description of the aircraft's areas, hereinafter, the forward-most part of the aircraft is assigned a relative longitudinal coordinate of 0, and the aft-most part of the aircraft is assigned a relative longitudinal coordinate of 100%. In this convention, the forward-most part is determined to correspond to a longitudinal range of 0 to 40% of the overall length of the aircraft, and the aft-most part is determined to correspond to a longitudinal range of 60 to 100% of the overall length of the aircraft. Furthermore, it is advantageous if the two propulsion devices located in the forward region are on a common straight line oriented parallel to the transverse direction. It is also appropriate if the two propulsion devices arranged in the rear area are on a common straight line aligned parallel to the transverse direction.

[0032] Advantageously, the propulsion devices in the forward area are arranged along a first straight line extending parallel to the transverse direction, and the propulsion devices in the rear area are arranged along a second straight line extending parallel to the transverse direction, so that the center of gravity of the aircraft when performing hovering flight is located at a longitudinal distance l1 from the straight line along which the propulsion devices in the forward area are arranged;

number

number

[0033] Preferably, the aircraft is further designed to have associated axes of rotation aligned parallel during hover flight.

[0034] Finally, it should be noted that according to the present invention, it is not excluded that the aircraft also includes, in addition to at least two propulsion devices that contribute to the effects according to the present invention, other propulsion devices that do not rotate in substantially the same direction of rotation.

[0035] According to a second aspect of the present invention, there is provided an aircraft including an aircraft fuselage and at least three propulsion devices mounted about the aircraft fuselage, rotatable about respective associated axes of rotation, and generating respective associated thrust vectors, the aircraft adapted to perform hover flight whereby all forces acting on the aircraft and all torques acting on the aircraft substantially vanish relative to the aircraft's center of gravity, wherein during hover flight, the associated axes of rotation of two of the at least three propulsion devices are substantially aligned in a first direction and the associated axes of rotation of other of the at least three propulsion devices are substantially oriented in a second direction, the first direction being non-parallel to the second direction, and wherein each of the two propulsion devices whose rotation axes are oriented in the first direction rotate in substantially the same rotational direction about their respective associated axes of rotation during hover flight.

[0036] For the understanding of the terms "substantially aligned in a first / second direction" and "rotating in substantially the same direction" according to the present invention, reference is made to the first aspect of the present invention, and the definitions set out therein apply accordingly to the second aspect.

[0037] If a (reference) axis facing a first direction is not parallel to a (reference) axis facing a second direction, the first direction is not parallel to the second direction. Preferably, the angle between the first direction and the second direction is in the range of 30° to 110°, preferably in the range of 40° to 100°, more preferably in the range of 60° to 95°.

[0038] Preferably, the at least three propulsion devices are mounted substantially in one plane around the aircraft fuselage. Preferably, the aircraft fuselage lies in a plane, i.e. a plane intersecting the aircraft fuselage. Furthermore, it is advantageous if the first direction and the second direction lie in this plane.

[0039] Here, "mounted substantially in a plane" does not mean that the propulsion devices or their mounting points need not be contained exactly in a plane. Therefore, it is within the scope of the present invention if one or more of the propulsion devices are pivotally mounted out of plane and / or are offset vertically relative to the plane. Conveniently, the vertical offset is constrained by the vertical extent of the aircraft fuselage, i.e., conveniently, the propulsion devices are mounted such that their axes of rotation are contained within a region of space formed between two horizontal planes that are in contact with the aircraft fuselage and spaced apart by the vertical extent of the aircraft fuselage. The vertical extent is relative to the direction of gravity when the aircraft is resting on (flat) ground.

[0040] Preferably, each of the rotation axes of two of the at least three propulsion devices oriented substantially in the first direction is oriented to be substantially parallel to a line passing through the two propulsion devices, which line is advantageously passed through the geometric center (this term is further explained below) or bearing point of the propulsion device.

[0041] According to the invention, an axis of rotation is substantially parallel to a line if the angle included between the axis of rotation and the line is less than 45°, preferably less than 30°, more preferably less than 15°.

[0042] Particularly preferably, the aircraft according to the second aspect of the invention includes at least four propulsion devices mounted about the aircraft fuselage, rotatable about respective associated axes of rotation, and generating respective associated thrust vectors, the aircraft thereby being adapted to perform hovering flight, during hovering flight the associated axes of rotation of two of the at least four propulsion devices are oriented substantially in a first direction and the associated axes of rotation of another two of the at least four propulsion devices are oriented substantially in a second direction, each of the two propulsion devices having axes oriented in the first direction rotate in substantially the same rotational sense about their respective associated axes of rotation during hovering flight, and / or each of the two propulsion devices having axes oriented in the second direction rotate in substantially the same direction about their respective associated axes of rotation during hovering flight.

[0043] The advantages of the aircraft according to the second aspect of the invention compared to the prior art correspond in principle to those already described in relation to the aircraft according to the first aspect of the invention. Therefore, to avoid repetition, reference is first made to that description, which concerns the use of the positive contribution of the Magnus effect, in particular when the propulsion devices rotate in the same direction. In connection with the latter contribution of the Magnus effect, it must be taken into account that in the arrangement of the propulsion devices around the aircraft fuselage, also referred to below as "star-shaped", in forward flight, as a rule, only some of the propulsion devices are irradiated with air in the direction of flight. Therefore, the Magnus effect in forward flight is greatest for propulsion devices whose rotation axes are oriented essentially perpendicular to the direction of flight, with substantially equal rotational rotation. That is, in the propulsion device arrangement according to the second aspect of the present invention, it is sufficient for the aircraft to be configured such that, during hover, each of the two propulsion devices with rotational axes oriented in a first direction rotates in substantially the same rotational direction about its associated rotational axis, or, in the case of at least four propulsion devices, each of the two propulsion devices with rotational axes oriented in a second direction rotates in substantially the same rotational direction about its associated rotational axis. In this case, two propulsion devices that do not rotate in substantially the same direction can rotate in opposite directions. When these two propulsion devices rotate in opposite directions, the torques directly cancel each other. However, when the aircraft includes at least four propulsion devices, it is particularly advantageous for the aircraft to be configured such that, during hover, each of the two propulsion devices with rotational axes oriented in a first direction rotates in substantially the same rotational direction about its associated rotational axis, and each of the two propulsion devices with rotational axes oriented in a second direction rotates in substantially the same rotational direction about its associated rotational axis. This ensures that the aircraft utilizes the positive effects of the Magnus effect in both the first and second directions during forward flight, and therefore makes the aircraft more flexible and stable during changes in flight direction.

[0044] In a particularly preferred embodiment, the air vehicle is further adapted such that, during hover flight, when one or more of the propulsion devices generate a particular predetermined thrust vector associated therewith, the center of gravity of the air vehicle is positioned such that all forces acting on the air vehicle and all torques acting on the air vehicle substantially disappear relative to the center of gravity of the air vehicle. Associated with this instruction is the constraint that the center of gravity of the air vehicle must be within a range determined by one or more of the propulsion devices such that the air vehicle is capable of hovering when driven with maximum thrust or maximum thrust vector. In other words, if the center of gravity is within said range, the propulsion devices are capable of generating the appropriate thrust vector to enable the air vehicle to perform hover flight.

[0045] Preferably, each of the rotation axes of two of the at least four propulsion devices oriented substantially in the first direction is oriented substantially parallel to a line passing through the two propulsion devices. It is preferred if each of the rotation axes of two further of the at least four propulsion devices oriented substantially in the second direction is oriented substantially parallel to a line passing through the two further propulsion devices, which line preferably passes through the geometric center or bearing point of the propulsion device.

[0046] As in the first aspect of the invention, compensation for one or more torques generated by the propulsion devices rotating in substantially the same direction is performed according to the invention by positioning the center of gravity of the aircraft in such a way that all forces and torques acting on the aircraft substantially vanish relative to the center of gravity of the aircraft during hover flight, taking into account the respective thrust vectors associated with the propulsion devices and predetermined for the propulsion devices. To be able to achieve a stable flight attitude during hover flight as well as forward flight, balance of all forces and torques acting on the aircraft must be achieved. This calculation is performed using the momentum theorem and angular momentum torque relationships already described and explained in relation to the first aspect of the invention. Therefore, the statements there also apply here and are further explained below.

[0047] It is advantageous if the three propulsion devices are arranged around the aircraft fuselage to form the sides of a triangle, preferably an equilateral triangle. The aircraft fuselage is advantageously located at the geometric center of the triangle. A first direction is defined by a line on which two of the three propulsion devices are located, and a second direction is substantially perpendicular to the first direction. Furthermore, the rotation axes of each of the two propulsion devices located on a line pointing in the first direction include angles with said line ranging from 0° to 45°, expediently 0° to 30°. The geometric center corresponds to the average of all points within the triangle (i.e., average over the area of ​​triangles with a constant density). If the angle between the rotation axes and the line pointing in the first direction is selected to be 30°, the rotation axes of the propulsion devices point towards (or away from) the geometric center. However, the angle can also be selected differently for each propulsion device. It is advantageous if the line is arranged through the geometric center or the support point of the propulsion device.

[0048] It is advantageous if the n propulsion devices are arranged around the periphery of the aircraft fuselage so as to form the sides of a polygon with n sides, n>3, expediently the sides of a regular polygon with n sides, n>3. Expediently, the aircraft fuselage is located at the geometric center of the n-sided polygon. Here, the first direction is defined by a first line on which two of the n propulsion devices are located, and the second direction is defined by a second line on which two further of the n propulsion devices are located. The rotation axes of each of the two propulsion devices located on the first line pointing in the first direction include an angle with the first line in the range of 0° to 45°, expediently 0° to 30°, expediently 0° to 20°, particularly preferably 0° to 18°. The rotation axes of different propulsion devices may thereby include different angles with the first line.

[0049] It is also suitable if the rotation axes of each of the two propulsion devices lying on a second straight line pointing in the second direction subtend an angle with the second straight line in the range of 0° to 45°, expediently in the range of 0° to 30°, 0° to 20°, particularly preferably in the range of 0° to 18°. The rotation axes of the different propulsion devices can thereby subtend different angles with the second straight line.

[0050] If the angles are selected as above, the axis of rotation of the propulsion device can point towards (or away from) the geometric centre of the n-sided polygon.

[0051] Particularly preferably, this means that the aircraft includes 3, 4, 5, 6, 7, 8, etc. propulsion devices arranged around the periphery of the aircraft fuselage, such that the propulsion devices form the sides of an equilateral triangle, a square, a regular polygon with 5, 6, 7 sides, or a regular polygon with 8 sides, etc. Advantageously, the aircraft fuselage is arranged substantially at the center of the n-sided polygon, in which case the geometric center, rather than the center of gravity, of the n-sided polygon is meant, since according to the invention the center of gravity of the aircraft does not necessarily have to coincide with the geometric center (geometric center of gravity). The geometric center of the n-sided polygon is defined according to the geometric center of the triangle.

[0052] Advantageously, n=2j, j>1. Therefore, it is even more advantageous if the aircraft fuselage is located between two opposing propulsion devices that are respectively in the shape of a regular polygon having 2j sides. In this case, it is advantageous if the rotation axes associated with each of the two specific opposing propulsion devices point substantially in a direction defined by a line on which the two specific opposing propulsion devices are located. Furthermore, it is advantageous if the aircraft is adapted to perform hovering flight by rotating each of the two opposing propulsion devices around their associated rotation axes in substantially the same direction during hovering flight. Therefore, in this case, the j direction according to the present invention can be defined.

[0053] Advantageously, the angle between the first and second straight lines is in the range of 60° to 100°, preferably 60° to 90°, particularly preferably 70° to 90°, and particularly preferably 72° to 90°. As will be shown later, for regular polygons with (2j+1) sides (j>1), it is particularly advantageous to select the first and second straight lines (or the corresponding directions) so that the angle between the first and second straight lines is 90°·(1−1 / (2j+1)). Thus, for (any) polygon with (2j+1) sides, a particularly preferred range for the angle between the first and second straight lines is given by [90°·(1−1 / (2j+1));90°]. It is possible to implement a configuration in which the rotation axis of a propulsion device positioned along a first straight line is oriented toward (or away from) the geometric center of a polygon having (2j+1) sides when the angle between the rotation axis of the propulsion device positioned along the first straight line is in the range [0°; 90° / (2j+1)] and / or the angle between the rotation axis of the propulsion device positioned along the second straight line is in the range [0°; 90° / (2j+1)].

[0054] For a regular polygon having 2j sides (j>1), it is advantageous to select the first and second lines so that they enclose an angle of 90°-90° / (2j)·(2j mod 4). Thus, the first and second lines each pass through the geometric center of the polygon having 2j sides. Therefore, for (any) polygon having 2j sides, a particularly preferred range of the angle between the first and second lines is given by [90°-90° / j;90°].

[0055] If the first and second straight lines are determined such that the angle between them is in the range [60°; 90°], and the angle between the rotation axis of a propulsion device arranged along the first straight line and the first straight line is in the range [0°; 30°], and / or the angle between the rotation axis of a propulsion device arranged along the second straight line and the second straight line is in the range [0°; 30°], the propulsion devices may be arranged around the aircraft fuselage in the form of an (arbitrary) regular polygon with n sides (n>2) such that the rotation axis of the propulsion device points towards (or away from) the geometric center. If n>3 is considered, it is sufficient that the angle between the rotation axis of the propulsion device and the first or second straight line passing through it is in the range [0°; 18°].

[0056] Conveniently, the second direction is substantially perpendicular, more preferably perpendicular, to the first direction, and two of the at least four propulsion devices are arranged along the first direction and two other of the at least four propulsion devices are arranged along a second direction that is substantially perpendicular to the first direction. This is an example where the propulsion devices may be arranged on the sides of a square around the periphery of the aircraft fuselage.

[0057] Preferably, the center of gravity of the aircraft when performing hovering flight is at a distance l in the first direction from a line along which the propulsion devices are arranged in the second direction. 34 placed in

number

number

[0058] Preferably, the center of gravity of the aircraft when performing hovering flight is at a distance l in a second direction from a line along which the propulsion devices are arranged in the first direction. 12 placed in

number

number

number

number

[0059] It may be advantageous for the aircraft according to the first aspect and the aircraft according to the second aspect to perform hover flight with approximately the same associated predetermined thrust vector.

[0060] Similarly, it may be advantageous for either the first or second aspect of the aircraft to further include a displacement device for displacing the center of gravity of the aircraft. Conveniently, in this regard, the aircraft further includes a fuel tank for supplying fuel to the propulsion devices and / or a battery for supplying power to the propulsion devices, and the displacement device is adapted to reposition fuel from the fuel tank or the battery within the aircraft, thereby positioning the center of gravity so that the aircraft performs hover flight when one or more of the propulsion devices generate a respective associated predetermined thrust vector. Thus, the center of gravity of the aircraft can be dynamically moved. An advantage of this is that the center of gravity of the aircraft can be optimally adapted to a wide variety of flight attitudes. The movement of the center of gravity can be performed by the aircraft control system.

[0061] Preferably, the aircraft according to the first or second aspect includes thrust vector control for individually controlling the thrust vectors of the propulsion devices.

[0062] According to a third aspect of the invention, there is provided a method of manufacturing an aircraft according to the first aspect of the invention, comprising the following steps: - each of the associated axes of rotation is oriented substantially transverse to the aircraft body; - when each of the at least two propulsion devices rotates in substantially the same rotational direction about its associated axis of rotation; - positioning the center of gravity of the aircraft such that one or more of the propulsion devices each generate a particular predetermined thrust vector associated with the one or more of the propulsion devices to cause the aircraft to perform a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially vanish relative to the center of gravity of the aircraft. A method is provided that includes:

[0063] According to a fourth aspect of the present invention, there is provided a method of manufacturing an aircraft according to the second aspect of the present invention, comprising the steps of: - associated axes of rotation of two of the at least three propulsion devices are oriented substantially in a first direction and associated axes of rotation of other of the at least three propulsion devices are oriented substantially in a second direction; - when each of two propulsion devices having rotational axes aligned in a first direction rotates in substantially the same rotational direction about each associated rotational axis during hover flight; - positioning the center of gravity of the aircraft such that one or more of the propulsion devices each generate a particular predetermined thrust vector associated with the one or more of the propulsion devices to cause the aircraft to perform a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially vanish relative to the center of gravity of the aircraft. A method is provided, comprising:

[0064] In preferred cases where the aircraft includes at least four propulsion devices, the associated rotational axes of two of the at least four propulsion devices are oriented substantially in a first direction and the associated rotational axes of the other two of the at least four propulsion devices are oriented substantially in a second direction, and each of the two propulsion devices having rotational axes oriented in the first direction rotates in substantially the same rotational direction about their associated rotational axes during hover flight, and / or each of the two propulsion devices having rotational axes oriented in the second direction rotates in substantially the same rotational direction about their associated rotational axes during hover flight.

[0065] According to a fifth aspect of the present invention, there is provided a method of controlling an aircraft having an aircraft body defining a longitudinal direction, a vertical direction, and a transverse direction, the longitudinal direction corresponding to a tail-to-nose direction of the aircraft, the vertical direction corresponding to the direction of the Earth's gravitational force when the aircraft is resting on the ground, and the transverse direction being perpendicular to the longitudinal and vertical directions, and at least two propulsion devices each rotatable about an associated axis of rotation and generating an associated thrust vector, a first plurality of the propulsion devices being arranged along a first line parallel to the transverse direction and a second plurality of the propulsion devices being arranged along a second line parallel to the transverse direction, the first line being spaced from the second line, and the center of gravity of the aircraft being located between the first line and the second line with respect to the longitudinal direction, the method comprising the steps of: - determining thrust vectors associated with the at least two propulsion devices such that the aircraft will perform hover flight when each of the at least two propulsion devices has an associated rotational axis oriented substantially transverse to the aircraft body and when each of the at least two propulsion devices rotates in substantially the same rotational sense about its associated rotational axis; During hover flight, all forces acting on the aircraft and all torques acting on the aircraft relative to the aircraft's center of gravity substantially disappear; determining an associated thrust vector; - driving each of the propulsion devices in substantially the same direction of rotation such that each propulsion device generates a particular associated thrust vector; Includes:

[0066] According to a sixth aspect of the present invention, there is provided a method of controlling an aircraft having an aircraft fuselage and at least three propulsion devices supported around the aircraft fuselage, each rotatable about an associated axis of rotation and each generating an associated thrust vector, the method comprising: - determining thrust vectors associated with the aircraft to perform hover flight when two of the rotational axes associated with the at least three propulsion devices are oriented substantially in a first direction and rotate in substantially the same rotational sense about each associated rotational axis, and / or when others of the rotational axes associated with the at least three propulsion devices are oriented substantially in a second direction that is not parallel to the first direction, During hover flight, all forces acting on the aircraft and all torques acting on the aircraft relative to the aircraft's center of gravity substantially disappear; determining an associated thrust vector; - aligning associated rotational axes of two of the at least three propulsion devices substantially in a first direction and aligning associated rotational axis of another of the at least three propulsion devices substantially in a second direction; - driving each of the propulsion devices such that each propulsion device rotates in an associated rotational direction and generates a particular associated thrust vector; each of the propulsion devices having a rotational axis substantially oriented in the respective first direction rotates in substantially the same rotational direction about its associated rotational axis; a driving step; A method is provided, comprising:

[0067] Preferably, the method is for controlling an aircraft having at least four propulsion devices and comprises the following steps: - determining thrust vectors associated with the aircraft to perform hover flight when two of the rotational axes associated with the at least four propulsion devices are oriented substantially in a first direction and rotate in substantially the same rotational sense about each associated rotational axis, and / or when two other rotational axes associated with the at least four propulsion devices are oriented substantially in a second direction that is not parallel to the first direction and rotate in substantially the same rotational sense about each associated rotational axis, During hover flight, all forces acting on the aircraft and all torques acting on the aircraft relative to the aircraft's center of gravity substantially disappear; determining an associated thrust vector; - aligning associated rotational axes of two of the at least four propulsion devices substantially in a first direction and aligning associated rotational axes of two other of the at least four propulsion devices substantially in a second direction; - driving each of the propulsion devices such that each propulsion device rotates in an associated rotational direction and generates a particular associated thrust vector; Each of the propulsion devices having a rotational axis substantially oriented in the first direction rotates in substantially the same rotational direction about its associated rotational axis, and / or each of the two propulsion devices having a rotational axis substantially oriented in the second direction rotates in substantially the same rotational direction about its associated rotational axis. a driving step; Includes:

[0068] Preferably, in the method of controlling an aircraft according to the fifth or sixth aspect, all of the determined relevant thrust vectors are selected to be substantially identical.

[0069] Advantageously, the method of controlling an aircraft according to the fifth or sixth aspect comprises the following steps: - positioning the center of gravity of the aircraft such that, when the propulsion device generates a particular predetermined thrust vector associated with the propulsion device, all forces acting on the aircraft and all torques acting on the aircraft are substantially nullified relative to the center of gravity of the aircraft. Further includes:

[0070] The advantages of the methods according to the third to sixth aspects of the invention are the same as those already explained in relation to the aircraft according to the invention according to the first and second aspects, and therefore useful and advantageous preferred embodiments of the first and second aspects apply accordingly to the third to sixth aspects of the invention.

[0071] Preferably, in an aircraft or method according to any of the aspects of the present invention, each of the propulsion devices is structurally identical.

[0072] Particularly preferably, in any aircraft or method according to any aspect of the invention, the propulsion device comprises a cyclogyro rotor.

[0073] In a preferred embodiment, each cyclogyro rotor includes a plurality of rotor blades rotatable along a circular path about an associated rotation axis of each of the propulsion devices or cyclogyro rotors, and a pitch mechanism including a coupling device and an attachment device, each of the plurality of rotor blades being pivotally mounted by the attachment device about its rotor blade attachment axis parallel to the rotation axis of the propulsion device or cyclogyro rotor. The cyclogyro rotor also conveniently includes an offset device to which each rotor blade is coupled by the coupling device at a mooring point associated with each rotor blade. The offset device thereby defines an eccentric bearing axis mounted at an adjustable offset distance parallel to the rotation axis of the propulsion device or cyclogyro rotor. As a result, when the offset distance is set to a non-zero value, rotation of the rotor blades along a circular path about the rotation axis of the propulsion device or cyclogyro rotor causes pitch motion of the rotor blades.

[0074] However, in this case, since it is essentially gravity that must be countered, the lift requirements of the aircraft generally remain fairly constant and an increase is usually not necessary. However, with the aid of an offset device, this increase can again result in a reduction in thrust, resulting in a reduction in rotor power consumption.

[0075] Preferred embodiments of the present invention will now be described with reference to the following figures: [Brief explanation of the drawings]

[0076] [Figure 1]1 is a perspective view of an aircraft according to a first embodiment of the present invention; [Figure 2a] 1 is a schematic diagram of a propulsion device and the forces and torques acting thereon. [Figure 2b] 1 is a schematic diagram of a propulsion device of an aircraft in forward flight and the forces and torques acting thereon, taking into account the incident flow; FIG. [Figure 3a] 1 is a schematic top view of an aircraft according to a first aspect of the present invention; [Figure 3b] 1 is a schematic side view of an aircraft and the forces and torques acting thereon according to a first aspect of the present invention; FIG. [Figure 3c] 1 is an exemplary configuration of an aircraft having four parallel, equal-sized propulsion devices to illustrate the preferred center of gravity location of the aircraft. [Figure 4] 1 is a schematic top view of an aircraft according to a first embodiment of the present invention for generalizing the conditions for a stable flight attitude; FIG. [Figure 5] 1 is a perspective view of a propulsion device according to the present invention; [Figure 6] FIG. 2 is a perspective view of an aircraft according to a second embodiment of the present invention. [Figure 7a] FIG. 2 is a schematic top view of an aircraft and the forces and torques acting thereon according to a second aspect of the present invention; [Figure 7b] FIG. 2 is a schematic diagram of an aircraft configured according to a second aspect of the invention and showing the forces and torques acting thereon, viewed from a first side; [Figure 7c] FIG. 10 is a schematic view of an aircraft configured according to a second aspect of the invention and of the forces and torques acting thereon, viewed from a second side; [Figure 7d] 10 is an exemplary configuration of an aircraft according to a second aspect of the invention having four equally sized propulsion devices arranged in a star configuration to illustrate the preferred center of gravity location of the aircraft. [Figure 8a] FIG. 2 is a cross-sectional top view of an aircraft having n propulsion devices according to a second embodiment of the present invention to illustrate the determination of the center of gravity; [Figure 8b] FIG. 1 is a side cross-sectional view of an aircraft having n propulsion devices. [Figure 9a]FIG. 2 is a schematic diagram of an aircraft according to a second embodiment of the invention having three propulsion devices. [Figure 9b] FIG. 2 is a schematic diagram of an aircraft according to a second embodiment of the invention having seven propulsion devices. [Figure 9c] FIG. 2 is a schematic diagram of an aircraft according to a second embodiment of the invention having six propulsion devices. DETAILED DESCRIPTION OF THE INVENTION

[0077] 1 shows a perspective view of an aircraft 100 according to a first embodiment of the invention, comprising an aircraft body 120 and a plurality of propulsion devices 1F, 1R. Each propulsion device 1F, 1R may be attached to the aircraft body 120 by a respective mounting or support device.

[0078] The illustrated aircraft 100 may be, for example, an air vehicle, a manned aircraft, a drone, or a so-called micro air vehicle (MAV).

[0079] To further describe the aircraft, a coordinate system is introduced that defines a longitudinal direction 101 or longitudinal axis, a transverse direction 102 or transverse axis, and a vertical direction 103 or vertical axis. The coordinate system is considered to be rigidly fixed 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. Thus, in the exemplary embodiment shown in FIG. 1 , the longitudinal direction 101 lies 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) of the aircraft 100 to the bow 121 or nose 121 (i.e., the front) of the aircraft 100. The vertical direction 103 or vertical axis corresponds to the direction of the Earth's gravitational force when the aircraft 100 is resting on (flat) ground. In other words, the vertical direction 103 is perpendicular to the aforementioned horizontal plane that contains the longitudinal direction 101. The transverse direction 102 or transverse axis is perpendicular to both the longitudinal direction 101 and the vertical direction 103. In other words, the transverse direction 102 lies in the aforementioned horizontal plane that contains the longitudinal direction 101 and is perpendicular to the longitudinal direction 101.

[0080] The illustrated aircraft 100 has four propulsion devices 1F, 1R. The illustrated propulsion devices 1F, 1R are cyclogyro rotors. Accordingly, the aircraft 100 shown in FIG. 1 may be referred to as a cyclogyro. The propulsion devices are described in more detail in connection with FIG. 5. Each of these propulsion devices 1F, 1R is mounted to rotate about an associated rotation axis 5. Each propulsion device 1F, 1R includes a plurality of rotor blades 2 pivotally mounted about their longitudinal axes. This allows the tilt angle of the rotor blades 2 to be changed during rotation of the propulsion devices 1F, 1R. By controlling not only the tilt angle of the rotor blades 2 but also the rotational speed (hereinafter also referred to as speed of rotation) of the propulsion devices 1F, 1R, not only the magnitude but also the direction of the thrust force generated, or the thrust vector describing it, can be varied.

[0081] 1 it can be seen that two of the four propulsion devices 1F are located at the front (nose) of the aircraft 100, and the other two propulsion devices 1R are located at the rear (tail) of the aircraft 100. The front and rear of the aircraft are defined as follows: the length of the aircraft is measured in the longitudinal direction 101, with the forward-most part of the aircraft (i.e. the nose 121 of the aircraft 100) being assigned a relative longitudinal coordinate of 0 and the aft-most part 122 of the aircraft 100 being assigned a relative longitudinal coordinate of 100%. By this convention, the forward part or extent is determined to correspond to a (longitudinal) range of 0 to 40% of the aircraft's total length, and the aft part or extent is determined to correspond to a (longitudinal) range of 60 to 100% of the aircraft's total length.

[0082] The two front propulsion devices 1F lie on a common straight line parallel to the transverse direction 102 or transverse axis, and similarly the two rear propulsion devices 1R lie on a common straight line parallel to the transverse direction 102 or transverse axis. It should be noted that said line does not necessarily have to be a common rotation axis to which the propulsion devices are (rigidly) coupled. Each propulsion device 1F, 1R can rotate via its own associated rotation axis 5, and it is also possible to control each propulsion device 1 individually, in particular to control their rotation speed separately. Furthermore, according to the invention, it is not necessary for all propulsion devices 1F, 1R to be in the same horizontal plane. As shown in FIG. 1, it may be advantageous if the two rear propulsion devices 1R of the aircraft are positioned higher relative to the two front propulsion devices 1F. This has the advantage that the rear propulsion devices 1R receive better inflow and are less susceptible to turbulence and turbulence caused by the front propulsion devices 1F.

[0083] The rotation axes 5 associated with the propulsion devices 1F, 1R are aligned parallel to the transverse direction 102 in the embodiment of Fig. 1. However, according to the invention, it is not absolutely necessary that all rotation axes 5 are parallel to one another. According to the invention, it is sufficient if each associated rotation axis 5 is substantially aligned with the transverse direction 102 of the aircraft body 120. According to the invention, a rotation axis 5 is substantially oriented in the transverse direction 102 of the aircraft body 120 if the angle included between the rotation axis 5 and an axis extending in the transverse direction and intersecting the rotation axis 5 is less than 45°, preferably less than 30°, particularly preferably less than 15°. The designation "substantially aligned in the transverse direction" therefore does not exclude the rotation axes 5 being exactly parallel to one another.

[0084] The aircraft 100 according to the present invention is designed to perform hovering flight by rotating each of the four illustrated propulsion devices 1F, 1R in the same rotational sense about their associated rotational axes 5. The design constraints this imposes on the aircraft 100 will be explained in relation to further figures, in particular Figures 3a and 3b.

[0085] In the general case where the axes of rotation 5 are substantially oriented in the transverse direction 102 of the aircraft body 120, it is necessary according to the invention that each of the propulsion devices 1 rotates in substantially the same direction of rotation about its associated axis of rotation 5. As already explained in detail in the introduction, this is fulfilled if the scalar product of the vector of the angular velocity of a particular propulsion device 1F, 1R and a fixedly predetermined vector arbitrarily oriented in the transverse direction 102 has the same sign for all propulsion devices 1R, 1F.

[0086] Figure 2a shows the forces 7 and torques 8 acting on a propulsion device 1 rotating about a rotation axis 5 at a rotational speed. In Figure 2a only a front view of the propulsion device 1 is shown, which is a schematic diagram. In the illustrated case, it is assumed that no air flows into the propulsion device 1. In the illustrated case, the propulsion device 1 rotates clockwise. The vector of the angular velocity corresponding to this rotation therefore points into the blade plane (according to the right-hand rule).

[0087] The thrust vector F,7 acting on the propulsion device 1 is perpendicular to the rotation axis 5 of the propulsion device 1. If a cyclogyro rotor is used as the propulsion device 1, the thrust vector F,7 is generated by the periodic adjustment of the rotor blades of the cyclogyro rotor. An offset device of the cyclogyro rotor allows the periodic rotor blade adjustment to be changed, and thus the thrust vector to be rotated in the entire plane perpendicular to the rotation axis 5 of the cyclogyro rotor, and the magnitude of the thrust vector to be changed. For convenience, thrust vector control is used for this purpose.

[0088] In addition to the thrust vector F,7, the propulsion device 1 generates a torque M,8 about the axis of rotation 5, opposite to the direction of rotation 51. This torque M,8 about the axis of rotation 5 results from the aerodynamic forces (lift and drag) of the propulsion device 1 or their tangential components; in the case of a cyclogyro rotor, the aerodynamic forces are primarily due to the rotating rotor blades. Therefore, to maintain a constant rotational speed, the propulsion device 1 must generate a (driving) torque that counteracts the torque generated by the aerodynamic forces. However, to enable the propulsion device 1 to generate such a (driving) torque during the flight phase, another torque M,8 is required, which the aircraft body must apply (according to the law of action and reaction) to "support" the propulsion device 1 in the air. This latter torque M,8 is approximately equal in magnitude (neglecting dissipation effects) to the torque generated by the aerodynamic forces and is directed in the same direction as the torque generated by the aerodynamic forces in order to maintain a constant rotational speed against the aerodynamic forces. Since the torque generated by the aerodynamic forces is opposite to the direction 51 of rotation of the propulsion device 1, the torque M,8 applied by the aircraft body is also opposite to the direction 51 of rotation of the propulsion device 1. Assuming that the torque due to the aerodynamic forces and the torque due to the propulsion device are substantially equal in magnitude but directed in opposite directions, the net torque remaining due to the rotation of the propulsion device 1 is the torque M,8 applied by the aircraft body.

[0089] This torque M,8 is therefore equal to the drive torque of the propulsion device 1. The torque M,8 can therefore be directly related to the magnitude of the thrust vector F,7. The design constraints of the aircraft according to the invention already mentioned in connection with Figure 1 and further described with reference to Figures 3a and 3b can therefore be stated using the mathematical-physical relationship between the torque M,8 and the thrust vector F,7.

[0090] Mathematically (and physically), the relationship between the thrust or corresponding thrust vector F,7 and the (driving) torque M,8 can be described by following the general equation of a propeller. Although a classical propeller differs from a cyclogyro rotor due to the position of the rotor blades relative to the axis of rotation, the generation of thrust in both concepts is based on the targeted unidirectional displacement of air by the rotor blades. The equations used below are derived in the appendix to this document for the sake of completeness.

[0091] First, consider the power required to displace air. This power P air can be derived from the so-called momentum theory (see Appendix) and results in P air =F*V a , (1) where F is the magnitude of the thrust vector, V a is the total air velocity in the plane of the propulsion device, the plane passing through the axis of rotation of the propulsion device and perpendicular to the direction of the airflow and therefore to the thrust vector F.

[0092] This power is provided by the propulsion device 1. First, the power P propulsion In general, the following applies: P propulsion =M*ω, (2) where M is the magnitude of the (driving) torque M,8 and ω is the rotational speed of the propulsion device 1 (magnitude of the vector of angular velocity).

[0093] Two powers P that can be expressed using the efficiency η air and P propulsion The relationship between can be expressed as follows: P air =η*P propulsion (3)

[0094] Efficiency η is the ratio of the driving power P propulsionThe ratio between the rotational speed ω of the propulsion device 1 and the radius r,52 on the one hand, and the total flow velocity V on the other hand a is a dimensionless parameter of the propulsion device 1, here denoted H (in the case of a propeller this is usually called the "advance rate").

number

[0095] The relationship between the (driving) torque,M,8,and the thrust or thrust vector,F,7,can therefore be established by starting from equation (3) and inserting equations (1), (2) and (4).

number

[0096] This relationship in this case depends only on the main quantities H, r and η of the propulsion device 1. Therefore, the relationship between the (magnitude of) the (driving) torque M,8 and the thrust or thrust vector F,7 is a linear function with a general proportionality coefficient a M=a*F (6) It can be expressed as:

[0097] This relationship will be used further below.

[0098] FIG. 2b shows a schematic representation of a propulsion device 1 in forward flight. The direction of movement of the aircraft including the illustrated propulsion device 1 is indicated by arrow 110. The torque M,8, which corresponds to the driving torque of the propulsion device 1, has already been described in connection with FIG. 2a. The propulsion device 1 is shown to be counter-flown by air 9 from the outside. The air flow 9 changes the aerodynamic characteristics of the propulsion device 1 and, therefore, the characteristics of the thrust vector generated. When the aircraft, and thus the propulsion device 1, is in forward flight, air actively flows from the front towards the propulsion device 1.

[0099] As mentioned in the introduction, the changing properties of the propulsion device 1 can be largely explained by the Magnus effect, which states that a sphere rotating in a flow experiences a lateral force perpendicular to the direction of the flow.

[0100] The direction of the lateral force here depends on the direction of rotation 51 of the body of the propulsion device 1. The propulsion force (its vertical component F rotor ,71 shown in Figure 2b), the Magnus effect provides an additional thrust or additional contribution F to the thrust vector. magnus , 72 are generated in the vertical direction. This increases the total thrust, or so-called lift, acting in the vertical direction of the propulsion device 1. However, in general, the lift requirements of an aircraft are roughly constant, and in this case, an increase is usually not necessary, since it is essentially the gravitational force of the Earth that must be counteracted.

[0101] Significant contribution to thrust vector during forward flight magnus , 72, the thrust vector contribution F generated by the propulsion device 1 rotor , 71 can be reduced, which is associated with a reduction in the power consumption of the propulsion device 1. Briefly, the Magnus effect replaces part of the thrust of the propulsion device 1, thereby reducing the power requirements during forward flight compared to hover flight.

[0102] However, if the incident flow 9 remains constant and the propulsion device 1 rotates in the opposite direction, the additional lateral force F of the Magnus effect magnus ,72 is the thrust F rotor , 71, thus reducing the total thrust or increasing the power requirement for the same desired lift.

[0103] In an aircraft according to the invention, the positive effects of the Magnus effect described are exploited in that all propulsion devices rotate in the same rotational sense about their associated rotational axes during hover and forward flight of the aircraft. As explained in more detail above, in a typical arrangement in which the rotational axes are oriented substantially transverse to the aircraft body, the propulsion devices rotate in substantially the same rotational sense.

[0104] When the propulsion devices 1 rotate in substantially the same direction about their respective associated axes of rotation, the faster the aircraft flies in forward flight, the greater the lateral force F magnus , 72 contributes more to the lift. That is, and to achieve a stable attitude during forward flight of the aircraft, it is sufficient to configure the aircraft for hover flight, where the incident air speed 9 is generally lowest.

[0105] The conditions for a stable flight attitude (balance of all forces and torques acting on the aircraft) in forward flight as well as in hover flight have already been generally stated in the introduction, from which the design constraints for the aircraft according to the first aspect of the invention are derived in the following in connection with Figures 3a and 3b.

[0106] In Figure 3a, an aircraft 100 according to a first aspect of the invention is shown in a highly schematic plan view. In addition to the aircraft body 120 already described in relation to Figure 1, the propulsion devices 1F and 1R, their associated axes of rotation 5 and longitudinal and transverse directions 101 and 102, and the centre of gravity S,150 of the aircraft 100 are also visible. The location or arrangement of the centre of gravity S,150 is crucial for balancing torques in the same direction generated by the propulsion devices 1 rotating in substantially the same sense of rotation. This will be explained in more detail with reference to Figure 3b.

[0107] Figure 3b shows in a side view and very schematic illustration the aircraft according to the first embodiment of the invention shown in Figure 3a in top view. 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 Figure 3b, the four propulsion devices 1F and 1R are arranged in a horizontal plane. However, the following description also applies when not all the propulsion devices are located in a horizontal plane. The rotation axes associated with the propulsion devices 1F and 1R are parallel to each other and to the transverse direction (pointing in the plane of the seat).

[0108] According to the present invention, all four propulsion devices 1F, 1R rotate in the same direction of rotation 51 at an assigned rotational speed. In FIG. 3b, all propulsion devices 1F and 1R rotate clockwise. This means that all four propulsion devices rotate clockwise relative to the transverse direction (y-axis) shown in FIG. 3a. In other words, the scalar product of each vector of angular velocity associated with propulsion devices 1F, 1R and a unit vector in the transverse direction is positive. Regardless of the reference frame used, the propulsion devices can also be described as rotating such that the surface of the propulsion device that first encounters the incoming air during forward flight rotates against the direction of the Earth's gravitational force. When the propulsion devices rotate clockwise, the Magnus effect has a particularly positive effect. This applies to any number of propulsion devices.

[0109] As explained above, thrust vectors are generated by the rotation of each propulsion device 1F, 1R. In the notation according to Fig. 3b, the thrust vector generated jointly by the two propulsion devices 1F located in the front area is indicated by F1,701, and the thrust vector generated jointly by the two propulsion devices 1R located in the rear area is indicated by F2,702. Since all propulsion devices 1F and 1R rotate in the same direction of rotation 51, all resulting (driving) torques M1,81, M2,82 also act in the same direction. Here, M1,81 denotes the (driving) torque of both front propulsion devices 1F, and M2,82 denotes the (driving) torque of both rear propulsion devices 1R.

[0110] Here, the momentum theorem and the angular momentum balance theorem are set about the aircraft's center of gravity S, 150, and in the illustrated case, only the momentum theorem in the vertical direction 103 (z-axis) and the angular momentum balance theorem about the transverse direction (y-axis) are relevant, since this is the only place where forces or torques act.

[0111] Therefore, the conditions for stable hovering are: ΣF z =F s -F1-F2=0 (7) ΣM s,y =F1*l1-F2*l2-M1-M2=0 (8) is.

[0112] The thrust vectors F1 and F2 can be adjusted to satisfy two balance conditions. Conveniently, the thrust vectors are set by the thrust vector control. l1,131 and l2,132 denote the longitudinal distances of the forward and aft propulsion devices 1F, 1R, respectively, from the center of gravity S,150. F S ,160 represents the force of the weight of the entire aircraft.

[0113] However, it is also possible to use these two equilibrium conditions to determine the center of gravity of the aircraft such that the hovering conditions are met for a given thrust vector F1 and thrust vector F2.

[0114] The torques M1,81 and M2,82 shown in FIG. 3b correspond to the driving torques of the two propulsion devices 1F and the two propulsion devices 1R, respectively. There is a mathematical physical relationship between the magnitudes of the torques M1,81 and M2,82 of the corresponding propulsion devices 1F and 1R and the magnitudes of the thrust vectors F1,701 and F2,702, respectively. This is determined by the above equation (6). That is, the magnitudes of the torques M1,81 and M2,82 are proportional to the magnitudes of the generated thrust vectors F1,701 and F2,702, respectively. Therefore, the torques cannot be freely controlled.

[0115] As noted above in relation to equation (6), the proportionality factor a of any propulsion device is essentially dependent on the efficiency of the propulsion device, its angular velocity, and other key quantities of the propulsion device.

[0116] Each propulsion device may have a different proportionality coefficient a. However, the values ​​of a for different propulsion devices of the same design or size are typically of the same order of magnitude. For convenience, they are essentially the same.

[0117] According to equation (6), the magnitudes M1 and M2 of the torque M1,81 and torque M2,82 are respectively: M i =a i *F i ,i∈{1,2} can be written as follows:

[0118] This results in the torque equation (8) being: ΣM s,y =F1*l1-F2*l2-F1*a1-F2*a2=0 It will look like this.

[0119] Here, this equation can be converted into the ratio of the magnitudes F1 and F2 of the two thrust vectors F1,701 and thrust vector F2,702, respectively.

Number

[0120] Equation (9) can serve as a constitutive equation of the aircraft. Equation (9) initially contains three freely selectable quantities (from the set of F1, F2, l1, l2). However, for a stable flight attitude, since Equation (7) also needs to be considered, only two of the aforementioned four quantities can be freely selected.

[0121] Therefore, there are several ways to satisfy Equation (7) and Equation (9).

[0122] (i) In the design of the first case, it may be necessary for the aircraft to be designed symmetrically. That is, the front rotation axis 5, i.e., the rotation axis of the propulsion device 1F arranged in the front region of the aircraft, and the rear rotation axis 5, i.e., the rotation axis of the propulsion device 1R arranged in the rear region of the aircraft, are equidistant from the center of gravity S,150. In other words, the center of gravity S,150 is located midway between the front rotation axis 5 and the rear rotation axis 5 in the longitudinal direction. In this case, l1 = l2. Therefore, Equation (9) and

Number

[0123] Therefore, in this configuration, the center of gravity S,150 tends to move forward, and as a result, l1 < l2, and the required thrust vectors F1 and F2 of the propulsion device 1F and the propulsion device 1R continue to increase respectively.

[0124] (ii) In the second case design, it is particularly preferred that the propulsion devices 1F and 1R are structurally identical, i.e. they are structurally identical, e.g. have the same size, same span, same number of rotor blades, same diameter and / or generate similar or the same (maximum) thrust / thrust vectors. Thus, in this case, F1=F2 or F1≈F2.

[0125] Since F1=F2≡F, from equation (7), initially, F=F S / 2. Therefore, from equation (9), l1-a1=l2+a2 is obtained.

[0126] If the longitudinal distance between the front propulsion device 1F and the rear propulsion device 1R is l=l1+l2, then from the previous equation,

number

[0127] It is understood that the center of gravity S,150 of the aircraft moves longitudinally by (a1 + a2) / 2 from the center l / 2 between the forward rotation axis 5 of the forward propulsion device 1F and the rear rotation axis 5 of the rear propulsion device 1R towards the rear rotation axis 5 of the rear propulsion device 1R. Typically, in this case, a1 = a2 ≡ a.

[0128] Now, if the aircraft is configured with structurally identical and equal-sized propulsion devices 1F and 1R for each pair of propulsion devices 1F and 1R, respectively, and therefore with approximately equal magnitudes of thrust / thrust vectors F1,701 and F2,702, the center of gravity S,150 can therefore be optimally positioned so that the torques M1,81 and M2,82 generated by the propulsion devices are purely balanced by the position of the center of gravity S,150. The optimal position is determined by equations (10) and (11).

[0129] It should be noted that here and below, only the position and center of gravity of the propulsion device in the longitudinal direction 101 plays a role in the considerations. The mounting or positioning of the center of gravity of the propulsion device with respect to the transverse and vertical directions 103 is irrelevant here and is left to the discretion of the skilled person. However, two, preferably symmetrical, bearings or positionings in the transverse and vertical directions 103 are preferred.

[0130] (iii) According to the present invention, it is also possible to combine the design aspects of the first case (i) with the design aspects of the second case (ii), i.e., the center of gravity S,150 of the aircraft can be shifted from the center between the front and rear rotation axes of the propulsion devices 1F and 1R, respectively, so that the conditions (7) and (8) for stable hovering flight at certain predetermined and even different thrust vectors / thrusts of the individual propulsion devices are met.

[0131] In practical applications, it is not always possible to arrange the mass of the aircraft so that the overall center of gravity S,150 can be positioned exactly at the specified optimum position described in case design (i), (ii), or (iii); for example, in case design (i), l = l, and in case design (ii), l and l are given by equations (10) and (11). Therefore, we define below the range within which the center of gravity S,150 can be positioned so as to still support a torque balance with the thrust force / thrust vector F,701 and thrust vector F,702 of the pair of propulsion device 1F and propulsion device 1R, respectively. For this purpose, the pair of propulsion devices i is arranged such that it is possible to ... i,max Suppose we can generate F i,max is the optimal configuration F i,opt This means that the aircraft must have at least F to remain in stable hover flight. i,optThis is because, in the preferred case, each pair of propulsion devices also generates a thrust surplus that can be used, among other things, to shift the position of the center of gravity S,150 from its optimal position. i,max is the maximum thrust allowed by the thrust vector control of the propulsion device, and therefore always the optimal design thrust F i,opt It must be more than or equal to this.

[0132] Considering the impulse theorem according to equation (7), F 1,opt ≦F 1,max ⇒ F 2,opt ≧F 2,min ≡F s -F 1,max is obtained.

[0133] This results in a maximum allowable thrust vector ratio

number

[0134] therefore, F 2,opt ≦F 2,max ⇒ F 1,opt ≧F 1,min ≡F s -F 2,max and Therefore, the minimum allowable thrust vector ratio is

number

[0135] These thrust vector ratios F1 / F2 are also described by equation (9), which can be used to determine that the maximum allowable distance of the center of gravity S,150 from the forward axis of rotation 5 in the longitudinal direction is:

number

number

[0136] The center of gravity S,150 is

number

[0137] Figure 3c serves to illustrate the above-mentioned area in which the centre of gravity S,150 of the aircraft can be suitably positioned to implement the invention according to the first aspect. Figure 3c shows a schematic view of an aircraft including propulsion devices 1F, 1R arranged along two straight lines, each parallel to the transverse direction of the aircraft. As already explained in relation to Figures 3a and 3b, for convenience, the aircraft includes four propulsion devices 1F, 1R, two of which 1F are arranged in the front region and two of which 1R are arranged in the rear region. Furthermore, it is assumed here that the propulsion devices 1F, 1R are structurally identical (as in case design (ii)), and that a1 = a2 ≡ a.

[0138] First, we further assume that torque compensation is realized purely by the location of the center of gravity, S,150,. F 1,opt =F 2,opt =F opt ⇒ M 1,opt =M 2,opt =M opt applies. In the example aircraft considered here, the total weight force generated by the corresponding total mass is: F s =1000N and the ratio / proportionality coefficient is typically a=0.2m, The distance of the propulsion device in the longitudinal direction is l=l1+l2=2m It is defined as follows.

[0139] Based on these specifications, the optimal center of gravity position is calculated using equations (10) and (11).

number

[0140] The aircraft's overall center of gravity S,150 is located at l 1,opt = 1.2 m, the range in which the position of the center of gravity S,150 can be positioned so that torque compensation can be compensated by the thrust / thrust vector of the propulsion devices 1F, 1R is defined here. For this purpose, for convenience, the maximum allowable thrust that can be generated by all the propulsion devices arranged along a straight line controlled by thrust vector control is F i,max =550N It is defined as follows.

[0141] This designation determines the maximum and minimum allowable thrust vector ratios.

number

number

[0142] That is, in this example, the center of gravity is conveniently located 1.1 to 1.3 m longitudinally from the front axis of rotation of the corresponding front propulsion device 1F.

[0143] Figure 4 shows a further embodiment of an aircraft 100 according to the first aspect of the invention. This figure mainly serves to generalize the results obtained in connection with figures 3a, 3b and 3c for any number K>2 of propulsion devices 1. It has already been pointed out above that in the discussion according to the invention, this is mainly a question of positioning the propulsion devices 1 in the longitudinal direction. The propulsion devices can therefore be arranged, for example, at different heights in the vertical direction. The longitudinal direction is distinguished in figure 4 as the x-axis 101. The K propulsion devices of an aircraft are aligned along N>1 straight lines g i As mentioned above, the line serves only to illustrate the geometrical arrangement of the propulsion device 1, not the structural components of the aircraft 100. i (denoted by the subscript i, i=1,…,N) on i (i=1,...,N) propulsion devices 1 are arranged.

number

[0144] Furthermore, the line g with index i i All n placed above i The propulsion device 1 has a size

number

[0145] straight g i (along which the propulsion devices 1 are arranged) are aligned parallel to the transverse direction 102, it should be noted that, according to the invention, it is not absolutely necessary that the rotation axes 5 of all the propulsion devices 1 are aligned (mathematically exactly) parallel to each other or to the transverse direction 102. It is sufficient if the rotation axes 5 of the propulsion devices 1 are substantially aligned with the transverse direction 102, in the sense defined at the beginning, especially during hover flight. In FIG. 4 it is shown that the rotation axes 5 of some of the propulsion devices 1 are not aligned exactly parallel to the transverse direction 102. According to the invention, the propulsion devices 1 are nevertheless aligned along a line g parallel to the transverse direction 102. i Because their geometric centers are located on such a line g i In order to satisfy the condition of being arranged on a parallel straight line, the support point of the propulsion device 1 must be located on such a straight line g. i It is also possible for it to be located substantially above.

[0146] Line g with index i i are the coordinates x in the longitudinal direction 101 (x-axis), respectively. i (i=1,…,N), where, without limiting generality, x i -x i-1 Assume that the line g is >0. i The longitudinal position x of i is fixed but arbitrary.

[0147] The center of gravity S, 150 of the aircraft 100 is aligned with the coordinate X S 3a, 3b, and 3c, the distance l1 and the distance l2 from the center of gravity are considered. i Note that we use coordinates relative to the longitudinal direction 101 of the line g, which turns out to be more convenient here. iCoordinates of and their distance l from the centroid S,150 i The relationship between l i =|x i -X S | is.

[0148] Using the introduced notation, from equations (7) and (8), the conditions for stable hovering or forward flight are as follows:

number

[0149] Substituting equation (13), equation (14) gives:

number

number

[0150] Note the intermediate result: the thrust vector F i Given this, from equation (15), the coordinate X of the center of gravity S,150 S However, equation (13) provides another condition that must be met for a stable flight attitude. Therefore, all N thrust vectors F i are not given arbitrarily, but only N-1 are given. That is, the position X of the center of gravity S,150 in a stable flight posture, especially in hovering flight, S is determined given N-1 thrust vectors. The values ​​of the thrust vectors given can of course also be equal.

[0151] The distance of the center of gravity S, 150 from the forward-most straight line g1 in the longitudinal direction or the propulsion device closest to the bow 121 or nose 121 of the aircraft 100 in the longitudinal direction is

number

[0152] Here, first, the line g i The propulsion device 1 arranged on the i That is, F1 ≒ F2 ≒ ≒ F N-1 ≒F N Consider the case where ≡F. Therefore, the center of gravity S, 150, is affected by the torque M generated by the propulsion device 1. i is optimally positioned so that it is balanced purely by the position of the center of gravity S,150. The optimal position is determined by equations (13) and (15). From equation (13),

number

[0153] Therefore, from equation (15),

number

[0154] In this case, a i We can conveniently assume that a,i=1,…,N.

[0155] The (longitudinal, x) coordinate of the center of gravity S, 150 S The maximum allowable range of can be determined using equations (13), (14) and (15), similar to the discussion of FIG. 3b for the general case above.

[0156] Figure 5 shows an embodiment of a propulsion device that can be used in an aircraft according to the invention. Each of these propulsion devices 1 is mounted rotatably about a rotation axis. Each propulsion device 1 includes a plurality of rotor blades 2 mounted to pivot about their longitudinal axes. This allows the inclination angle of the rotor blades 2 to be changed during rotation of the propulsion device 1. By controlling the inclination angle of the rotor blades 2 as well as the rotational speed of the propulsion device 1, the magnitude and direction of the generated thrust vector can be varied.

[0157] FIG. 5 shows a perspective view of a propulsion device 1 according to the present invention. The propulsion device 1 has a cylindrical shape. The illustrated propulsion device 1 is a cyclogyro rotor. The propulsion device 1 includes five rotor blades 2, each with an associated pitch mechanism 3, an offset device 4, and a disk 11. Propulsion devices with a different number of rotor blades are also possible. The rotor blades 2 are rotatably mounted about the axis of rotation of the propulsion device 1. The offset device 4 defines an eccentric bearing axis that is eccentrically mounted relative to the axis of rotation of the propulsion device 1. In FIG. 5, the offset device is shown as an offset disk. The offset disk is mounted so as to be freely rotatable about the eccentric bearing axis. The eccentric bearing of the offset disk 4 represents the 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 rotation about the axis of rotation of the propulsion device 1. Each of the illustrated pitch mechanisms 3 includes a coupling device 31 and a bearing device 33. Each rotor blade 2 is pivotally supported by a corresponding bearing device 33. The rotor blade 2 is supported around an axis parallel to the rotation axis of the propulsion device 1. This axis is the rotor blade bearing axis 33. The rotor blade 2 can be supported with the help of a bearing means, for example, one or more pins, so-called main pins. The bearing means is preferably part of the bearing device 33. The rotor blade bearing axis 33 can pass through the center of gravity of the rotor blade 2. However, it is preferable that the rotor blade 2 is supported away from the center of gravity. The coupling device 31 of the pitch mechanism 3 couples the rotor blade 2 to the offset device 4 so that the rotor blade 2 performs a pitch movement when rotating around the rotation axis of the propulsion device 1, provided that the eccentric bearing axis does not coincide with the rotation axis of the propulsion device 1. One end of the coupling device 31 is coupled to the offset device 4 at a mooring point. The other end portion of the coupling device 31 is coupled to the rotor blade 2.

[0158] The offset disk 4 is mounted so as to be freely rotatable. The rotation axis of the offset disk 4 preferably runs parallel to the rotation axis of the propulsion device 1 at an offset distance. This results in an eccentric mounting of the offset disk 4 relative to the rotation axis of the propulsion device 1. This offset distance may be adjustable. An offset device 4 with an adjustable eccentricity may be realized, for example, by means of a planetary gear. If the offset distance is not zero, a pitch movement of the rotor blades 2 occurs.

[0159] The coupling device 31 is coupled to the rotor blade 2 at a coupling point 32. For this purpose, the coupling device 31 may comprise a coupling means. In the propulsion device 1 shown in FIG. 5, the coupling device 31 comprises a connecting rod ("connecting rod") and a pin, a so-called pitch link pin. The pin is a constructive embodiment of a coupling means according to the invention. In the exemplary embodiment shown in FIG. 5, the coupling means 31 is coupled to the rotor blade 2 at the coupling point 32 by using a connecting element 61 rather than by a direct connection to the rotor blade 2. One end of the connecting element 61 is thereby rigidly connected to the rotor blade 2. This connection is preferably made at the rotor blade attachment 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 indirectly introduced to the rotor blade 2 via the connecting element 61 via the coupling device by means of the connecting rod 31.

[0160] However, according to the invention, a direct coupling of the coupling device 31 to the rotor blade 2 is also possible.

[0161] Because the coupling device 31 of the pitch mechanism is mounted eccentrically relative to the axis of rotation of the propulsion device 1, when the rotor blade 2 rotates about the axis of rotation of the propulsion device 1, the coupling point 32 moves in an arc relative to the rotor blade bearing axis 33. This causes a pitch movement of the rotor blade 2, which is therefore a pendulum movement of the rotor blade 2 about the rotor blade bearing axis 33.

[0162] The diameter of the propulsion device 1 corresponds to twice the distance from the rotation axis of the propulsion device 1 to the rotor blade bearing axis 33 or bearing point. This diameter is related to the blade speed during rotation and therefore to the thrust generated. In an exemplary embodiment of the propulsion device 1 according to the invention, the diameter is in the range of 150 mm to 2000 mm, preferably 300 mm to 500 mm, particularly preferably 350 mm.

[0163] 5 further includes a disk 11. This disk 11 is designed to aerodynamically separate the rotor blades 2 from the remaining components of the propulsion device 1. Such a disk 11 is particularly advantageous when the propulsion device 1 is operated at high speeds.

[0164] The span of the propulsion device 1 is defined by the length of the rotor blades 2. The span of the propulsion device 1 is the (longitudinal) distance between two disks 11.

[0165] The span of one of the cyclogyro rotors that can be used according to the invention is expediently between a few centimeters and 2 meters, preferably between 350 and 420 mm. Advantageously, several cyclogyro rotors are used in an aircraft according to the invention. Their span widths preferably differ from one another by a maximum of 25%, expediently by a maximum of 10%. Their diameters preferably differ from one another by a maximum of 25%, preferably by a maximum of 10%.

[0166] The rotor blades 2 shown in Figure 5 have a symmetrical profile. The present invention is not limited to propulsion devices having rotor blades with a symmetrical profile.

[0167] The propulsion device 1 generates thrust or a thrust vector through two coupled rotational motions. The first rotational motion is the rotation of the rotor blades 2 around the rotation axis of the propulsion device 1. This first rotational motion causes the rotor blades 2 to move along a circular path around the rotation axis of the propulsion device. Specifically, the rotor blade bearing axes 33 or rotor blade bearing points move along the circular path. Each rotor blade bearing axis 33 is parallel to the longitudinal axis of the rotor blade 2. The longitudinal axes of the rotor blades 2 are parallel to the rotation axis of the propulsion device 1. Therefore, the longitudinal axes of the rotor blades 2 are also parallel to the rotor blade bearing axes 33. The thrust direction of the propulsion device 1 is perpendicular to the rotation axis of the propulsion device 1. For optimal thrust generation, all rotor blades 2 should always be oriented as best possible with respect to the flow direction. This ensures that each rotor blade 2 contributes the greatest amount of thrust to the total thrust. As the propulsion device 1 rotates around its rotation axis, the pitch of each rotor blade 2 is continuously changed by the pitch mechanism described above. Each rotor blade 2 periodically changes its pitch angle, i.e., performs a reciprocating motion. This is the pitch motion. At the same time, the coupling point 32 moves on an arc around the rotor blade support axis 33. This is the second rotational motion.

[0168] The magnitude and direction of the generated thrust force / thrust vector depend on the pitch of the rotor blades 2. Therefore, the distance of the eccentric bearing of the offset device 4 or pitch mechanism 3 relative to the axis of rotation of the propulsion device 1 affects the magnitude of the generated thrust force / thrust vector. By moving the eccentric bearing of the offset device 4 circumferentially, i.e., by a fixed distance relative to the axis of rotation of the propulsion device 1, the direction of the generated thrust vector is changed.

[0169] 5, the pitch mechanism 3 is shown only on one side of the propulsion device 1, but for stability reasons it may be appropriate to also mount a corresponding pitch mechanism on the other side of the propulsion device. For example, a pitch mechanism could also be mounted in the middle of the propulsion device.

[0170] 6 shows a perspective view of an aircraft 200 according to a second embodiment of the present invention, including an aircraft fuselage 220 and a plurality of propulsion devices 1A and 1B. Four propulsion devices 1A and 1B are shown arranged around the periphery of the aircraft fuselage 220. Each of the propulsion devices 1A and 1B is connected to the aircraft fuselage 220 via an arm 221 and an arm 222, respectively. The propulsion devices 1A and 1B may be attached to the arms 221 and 222, respectively, by a suitable mounting or support device. The presence of the arms 221 and 222, respectively, is not essential. The propulsion devices 1A and 1B may also be coupled to the aircraft body 220 in other ways. Here, the aircraft body 220 and the propulsion devices 1A and 1B are substantially in one plane.

[0171] The illustrated aircraft 200 may be, for example, an air vehicle, a manned aircraft, a drone, or a so-called micro air vehicle (MAV).

[0172] To further describe the illustrated aircraft 200, a reference frame is introduced that defines a first direction 201, a second direction 202, and a vertical direction 203 or axis. The vertical direction 203 or axis corresponds to the direction of the Earth's gravitational force when the aircraft 200 is resting on the ground. The vertical direction 203 is perpendicular to the aforementioned 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 their respective axes, lie within the plane and are therefore each perpendicular to the vertical direction. It is essential for the aircraft 200 of the second aspect of the invention considered herein that the first direction 201 and the second direction 202 are not parallel to each other. In the illustrated embodiment, the first direction 201 and the second direction 202 are perpendicular to each other.

[0173] The direction thus defined is rigidly fixed to the aircraft 200.

[0174] The illustrated aircraft 200 has four propulsion devices 1A and 1B. The illustrated propulsion devices 1A and 1B are each cyclogyro rotors. A more detailed description of cyclogyro rotors has already been provided in connection with FIG. 5 . Each propulsion device 1A and 1B is mounted to rotate about an associated rotation axis 5. Each propulsion device 1A and 1B includes a plurality of rotor blades 2 mounted to pivot about their longitudinal axes. This allows the tilt angle of the rotor blades 2 to be varied during rotation of the propulsion device 1A or 1B. By controlling not only the tilt angle of the rotor blades 2 but also the rotational speed (hereinafter also referred to as speed of rotation) of each propulsion device 1A and 1B, not only the magnitude but also the direction of the generated thrust and the thrust vector describing it can be varied, respectively.

[0175] 6, it can be seen that the four propulsion devices 1A and 1B essentially form the sides of a rectangle or square. The fuselage 220 is located at the geometric center of this rectangle or square. For convenience, the propulsion devices 1A and 1B are each equidistant from the center or fuselage. For this purpose, the arms 221 and 222 may have the same length. In this case, the propulsion devices 1A and 1B are located on the sides of the square.

[0176] Two propulsion devices 1A corresponding to opposite sides of the rectangle or square lie on a common straight line, which in the illustrated example is substantially parallel to a first direction 201, and similarly, two propulsion devices 1B corresponding to opposite sides of the rectangle or square lie on a common straight line substantially parallel to a second direction 202. It should be noted that said straight line does not necessarily have to be a common axis of rotation to which the propulsion devices are (rigidly) coupled. Each propulsion device 1A, 1B can rotate via its own associated axis of rotation 5A, 5B, and it is also possible that each propulsion device 1A, 1B can be controlled individually, in particular to control their rotational speed separately.

[0177] In the embodiment of FIG. 6 , the rotation axis 5A associated with the propulsion device 1A is substantially aligned with the first direction 201. In the example embodiment of FIG. 6 , the rotation axis 5B associated with the propulsion device 1B is substantially aligned with the first direction 202. It can be seen in FIG. 6 that the rotation axes 5A, 5B are not aligned exactly parallel to the first direction 201 or the second direction 202. In fact, according to the present invention, this is already the case when each of the associated rotation axes 5A, 5B is substantially aligned with the first direction 201 and the second direction 202, respectively. According to the present invention, the rotation axis 5A is substantially aligned with the first direction 201 if the angle included between the rotation axis 5A and an axis extending in the first direction 201 and intersecting the rotation axis 5A is less than 45°, preferably less than 30°, more preferably less than 15°. Therefore, the designation "substantially aligned with the first direction" does not exclude the rotation axis 5A being exactly parallel to the first direction 201. The same applies to the axis of rotation 5B and the second direction 202 of the second propulsion device 1B.

[0178] The aircraft 200 according to the present invention is designed to perform hover flight by rotating each of the two illustrated propulsion devices 1A in substantially the same rotational sense about its associated axis of rotation 5A and / or by rotating each of the two illustrated propulsion devices 1B in substantially the same rotational sense about its associated axis of rotation 5B, which imposes design constraints on the aircraft 200 as will be explained in relation to further figures, particularly Figures 7a and 7b.

[0179] In Figure 7a an aircraft 200 according to a second aspect of the invention is shown in a highly schematic plan view. Firstly, the aircraft fuselage 220 already described in relation to Figure 6, the propulsion devices 1A1, 1A2 and 1B3, 1B4, their associated axes of rotation 5A and 5B, and the first and second directions 201 and 202 can be seen, with the first direction 201 being perpendicular to the second direction 202.

[0180] To explain the mathematical and physical relationships, it is useful to introduce a (Cartesian) Cartesian coordinate system. In Figures 7a and 7b, a Cartesian coordinate system is used with an x-axis, a y-axis and a z-axis. It should be noted that, in general, the first and second directions according to the invention do not necessarily 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, and the (Cartesian) Cartesian coordinate system is intended to serve as an intentional mathematical description of the aircraft.

[0181] Also shown is the center of gravity S,250 of the aircraft 200. The location or placement of the center of gravity S,250 is important for balancing torques in the same direction generated by propulsion devices 1A1, 1A2 and 1B3, 1B4, which rotate in substantially the same direction, as will be explained in more detail with respect to FIG. 7b. In the illustrated example, the center of gravity S,250 is positioned such that the aircraft 200 can utilize the Magnus effect in both forward flight in a (positive) first direction 201 (here, coinciding with the positive x-direction) and in a (positive) second direction 202 (here, coinciding with the positive y-direction). When the aircraft 200 moves in the first direction 201 in forward flight, the propulsion devices 1B3, 1B4 rotate in substantially the same rotational direction, advantageously clockwise, about their associated axes of rotation 5B. As defined above in connection with the first embodiment, this means that the two propulsion devices 1B3, 1B4 rotate clockwise with respect to the second direction (y-axis) shown in FIG. 7a. In other words, the scalar product of each vector of angular velocity associated with the propulsion devices 1B3, 1B4 and the unit vector of the second direction is positive. Regardless of the reference frame used, the propulsion devices 1B3, 1B4 can also be described as rotating such that the surface of the propulsion devices 1B3, 1B4 that first encounters the incoming air in forward flight rotates in a direction opposite to the direction of the Earth's gravitational force. When the aircraft 200 moves in the second direction 202 in forward flight, the propulsion devices 1A1, 1A2 rotate in substantially the same rotational direction about the associated rotation axis 5A, advantageously counterclockwise. The above definitions apply accordingly. In the coordinate system shown in FIG. 7a, this means that the scalar product of each vector of angular velocity associated with the propulsion devices 1A1, 1A2 and the unit vector of the first direction is negative. Regardless of the reference frame used, it is also true that the propulsion devices 1A1, 1A2 rotate such that the surface of the propulsion devices 1A1, 1A2 that first encounters the incoming air during forward flight rotates against the direction of the Earth's gravitational force.

[0182] Finally, thrust vectors F1, 2001, F2, 2002, F3, 2003, and F4, 2004 are depicted, which are generated by the rotation of the propulsion devices about rotation axis 5A and rotation axis 5B, respectively. Thrust vectors F1, 2001, F2, 2002, F3, 2003, and F4, 2004 point out of the image plane, i.e., lift forces are generated.

[0183] In forward flight in the first direction (x-axis), it is also possible for propulsion devices 1A1 and 1A2 to rotate in opposite directions, one clockwise and the other counterclockwise, while propulsion devices 1B3 and 1B4 rotate in the same direction. The same applies to forward flight in the second direction (y-axis). The directions of thrust vectors F1, 2001, F2, 2002, F3, 2003, and F4, 2004 remain unaffected.

[0184] Figures 7b and 7c show different side views and highly schematic illustrations of the aircraft according to the second embodiment of the invention shown in the top view of Figure 7a. In the side view of Figure 7b, two propulsion devices 1A1, 1A2 and one of two propulsion devices 1B3, 1B4 are visible. In the side view of Figure 7c, two propulsion devices 1B3, 1B4 and one of two propulsion devices 1A1, 1A2 are visible. The rotation axis 5A associated with the propulsion devices 1A1, 1A2 is parallel to a first direction 201 (here, the x-direction), and the rotation axis 5B associated with the propulsion devices 1B3, 1B4 is parallel to a second direction (here, the y-direction) (pointing in the plane of the seat).

[0185] In the embodiment considered by the present invention, the propulsion devices 1B3, 1B4 rotate in the same direction of rotation 251 at an assigned rotational speed. In Fig. 7b, the two propulsion devices 1B3, 1B4 rotate clockwise as defined above. As already explained, the rotation of each propulsion device 1B3, 1B4 generates a thrust vector. In the representation according to Fig. 7b, the thrust vector generated together by the two propulsion devices 1B3, 1B4 is F 34 , 2034, where F 34= F3 + F4 (see FIG. 7a). Since the propulsion devices 1B3 and 1B4 rotate in the same direction of rotation 251, the total (driving) torque M 34 ,280 also acts in the same direction, where M 34 , 280 indicates the (driving) torque of both propulsion devices 1B3, 1B4, i.e., M 34 =M3+M4.

[0186] The propulsion devices 1A1, 1A2 generate thrust vectors F1, 2001 and F2, 2002, respectively. The direction of rotation of the propulsion devices 1A1, 1A2 is not important in the present discussion, which considers designing an aircraft that favors forward flight in a first direction 201. However, for reasons of symmetry, it is preferable to design the aircraft such that a stable flight attitude, and in particular stable forward flight, is possible even with equally rotating propulsion devices 1A1, 1A2, as will be explained further below.

[0187] 7b, the momentum theorem and the angular momentum balance theorem are applied with respect to the aircraft's center of gravity S, 250, and in the illustrated case, only the momentum theorem in the vertical direction 203 (z-axis) and the angular momentum balance theorem about the second direction 202 (y-axis) are relevant, since this is the only place where forces or torques act.

[0188] Therefore, the conditions for stable hover flight are: ΣF z =F s -F1-F2-F 34 =0 (17) ΣM s,y =F1*l1+F 34 *l 34 -F2*l2-M 34 =0 (18) is.

[0189] Thrust vectors F1, F2 and F 34 The magnitude of the force can be adjusted to satisfy two equilibrium conditions. It is convenient to adjust the thrust vector by thrust vector control.

[0190] However, with two equilibrium conditions, for certain thrust vectors F1, F2, and F 34 It is also possible to determine the center of gravity of the aircraft so that the hovering conditions for

[0191] The torque M shown in Fig. 7b 34 , 280 corresponds to the (driving) torque of both propulsion devices 1B3, 1B4. As already explained in connection with the first aspect of the invention, the torque M 34 ,280 magnitude and thrust vector F 34 There is a mathematical physical relationship between the magnitude of a and the coefficient of proportionality a, which is determined by equation (6) above. Each propulsion device may have a different proportionality coefficient a. However, the values ​​of a for different propulsion devices of the same design or size are typically of the same order of magnitude. For convenience, they are essentially the same.

[0192] According to equation (6), the magnitudes of torques M1, M2, M3, and M4 are: M i =a i *F i ,i∈{1,2,3,4} can be written as follows:

[0193] In the embodiment considered, the rotation axes of the propulsion devices 1B3, 1B4, which have the same direction of rotation, are aligned in parallel, so that M3 and M4 are parallel, and therefore the same applies in terms of size; |M3+M4|=M3+M4=M 34 ≡a 34 *F 34 is.

[0194] It should be noted here that the above equations also serve as good approximations for the commonly considered case of essentially co-directional rotating axes.

[0195] This results in the torque equation (18) ΣM s,y =F1*l1+F34 *l 34 -F2*l2-F 34 *a 34 =0 is obtained.

[0196] Here, F1 and F2 indicate the magnitudes of the thrust vectors F1,2001; F2,2002 generated by the propulsion devices 1A1 and 1A2, respectively, and l1,231 is the distance of the thrust vector F1,2001 from the center of gravity S,250 of the aircraft determined in the first direction (this distance l1 can be identified as the distance in the first direction between the center of gravity S,250 of the aircraft and the geometric center along the rotation axis 5A of the propulsion device 1A1; in other words, l1 is the distance of the thrust vector F1,2001 from the center of gravity S,250 of the aircraft). l2 is the distance in the first direction from the center of gravity S,250 of the aircraft to half the span of the propulsion device 1A2), l2 is the distance of the thrust vector F2,2002 from the center of gravity S,250 of the aircraft determined in the first direction (this distance l2 can be identified as the distance in the first direction between the center of gravity S,250 of the aircraft and the geometric center of the propulsion device 1A2 along the rotation axis 5A; in other words, l2 is the distance in the first direction from the center of gravity S,250 of the aircraft to half the span of the propulsion device 1A2), F 34 is the magnitude F of the thrust vector generated by both the propulsion device 1B3 and the propulsion device 1B4. 34 =F3+F4,2034, and l 34 , 234 is determined with respect to the first direction, the center of gravity S, 250 of the aircraft, and on the other hand, the thrust vector F 34 , 2034, or, on the other hand, equivalently, the rotation axes of the propulsion devices 1B3 and 1B4, or, equivalently, the distance between the straight line passing through the propulsion devices 1B3 and 1B4 (assuming here that the propulsion devices 1B3 and 1B4 are on a straight line at least approximately parallel to the second direction), a 34 is a proportionality coefficient assigned to the propulsion device 1B3 and the propulsion device 1B4.

[0197] Here, this equation can be converted into the ratio of the magnitude F1 to the magnitude F2 of the two thrust vectors F1,2001 and F2,2002, respectively.

number

[0198] Equation (19) can serve as the constitutive equation for the aircraft. Equation (19) is initially written as (F1, F2, F 34 , l1, l2, l 34 However, in a stable flight attitude, only three of the four aforementioned quantities can be freely selected, since equation (17) must also be taken into account.

[0199] The corresponding constitutive equations are also obtained for the case where the momentum theorem is established in the vertical direction 203 (z-axis) and the angular momentum balance theorem is established around the first direction 201 (x-axis). For this purpose, reference is made to Figure 7c. Such considerations are necessary if we wish to use the effects according to the invention also for forward flight in the second direction (y-axis), i.e., in particular the positive contribution of the Magnus effect.

[0200] Therefore, the conditions for stable hover flight are: ΣF z =F s -F3-F4-F 12 =0 (20) ΣM s,x =F3*l3+F 12 *l 12 -F4*l4-M 12 =0 (21) is.

[0201] These terms are the same as in equations (17) and (18), but the subscripts change from 1 to 3, 2 to 4, 3 to 1, and 4 to 2. Therefore, the individual equations will not be repeated. In particular, M 12 , 285 is the total torque generated by the propulsion devices 1A1, 1A2.

[0202] Considering the related observations of equations (17) and (18) with respect to equation (6), the torque equation (21) becomes: ΣM s,x =F3*l3+F 12 *l 12 -F4*l4-F 12 *a 12 =0 can be written as follows:

[0203] Here, F3,F4 indicate the magnitudes of the thrust vectors F3,2003; F4,2004 generated by the propulsion device 1B3 and the propulsion device 1B4, respectively (see FIG. 7a), and l3,236 is the distance of the thrust vector F3 from the center of gravity S,250 of the aircraft determined in the second direction (this distance l3 can be identified as the distance in the second direction between the center of gravity S,250 of the aircraft and the geometric center of the propulsion device 1B3 along the rotation axis 5B; in other words, l3 is the distance from the center of gravity S,250 of the aircraft to the geometric center of the propulsion device 1B3 along the rotation axis 5B). l4,237 is the distance from the center of gravity S,250 of the aircraft to half the span of the propulsion device 1B3 in the second direction), l4,237 is the distance of the thrust vector F4 from the center of gravity S,250 of the aircraft, determined in the second direction (this distance l4 can be identified as the distance in the second direction between the center of gravity S,250 of the aircraft and the geometric center of the propulsion device 1B4 along the axis of rotation 5B; in other words, l4 is the distance in the second direction from the center of gravity S,250 of the aircraft to half the span of the propulsion device 1B4), F 12 is the magnitude F of the thrust vector generated by both the propulsion device 1A1 and the propulsion device 1A2. 12 =F1+F2,2012, and l 12 ,239 is determined with respect to the second direction, the center of gravity S of the aircraft, 250, and on the other hand the thrust vector F 12 , 2012, or, on the other hand, equivalently, the rotation axes of the propulsion devices 1A1 and 1A2, or, equivalently, the distance between the propulsion devices 1A1 and 1A2 (here, it is assumed that the propulsion devices 1A1 and 1A2 are on a line at least approximately parallel to the first direction), a 12is a proportionality coefficient assigned to the propulsion device 1A1 and the propulsion device 1A2.

[0204] Now, this equation can be converted into the ratio of the magnitude F3 to the magnitude F4 of the two thrust vectors F3 and F4, respectively.

number

[0205] Due to the star topology of the thrust devices 1A1, 1A2, 1B3 and 1B4, it is convenient if in each case a pair of thrust devices 1A1, 1A2 or thrust devices 1B3, 1B4 generates half of the required thrust. This allows the boundary conditions F 12 =F 34 (twenty three) is brought about.

[0206] Note that this does not necessarily imply that all thrust vectors F1, F2, F3, and F4 must be the same; it is sufficient if the sum of the thrust vectors of the two opposing propulsion devices is the same. However, all thrust vectors F1, F2, F3, and F4 can also be individually different.

[0207] Another useful boundary condition arises when the propulsion device is required to be mounted, preferably at the center, of the aircraft fuselage 220, namely:

number

[0208] From these boundary conditions (23) and (24a), the following constitutive equation is obtained:

number

number

[0209] The next step is now to determine the number of thrust vectors that can be freely preset using equations (17), (18), (20) and (21). Assuming that the positions of the propulsion devices are fixed, the equations contain the following unknowns: F1, F2, F3, F4, l 12 and l 34 Furthermore, it should be noted that equation (17) and equation (20) give the same constraints. Therefore, we have three equations for six unknowns. The center of gravity is l 12 and l 34 and equations (17), (18), (20), and (21) specify the thrust vectors. Three of the four thrust vectors, F1, F2, and F3, can be specified as desired. If additional boundary conditions are considered, the number of freely definable thrust vectors decreases accordingly.

[0210] There are several ways to satisfy equations (17), (20), (25a), and (25b).

[0211] (i) In the first case, the design may require that the aircraft be designed symmetrically, i.e., the center of gravity S,250 is exactly halfway between the propulsion devices 1A1, 1A2 and / or 1B3, 1B4. In this case, l1 = l2 and / or l3 = l4. From equations (25a) and (25b),

number

[0212]

Number

[0213] Therefore, in the design of this case, the center of gravity S, 250 tends to move in the positive first direction and / or the second direction. As a result, l1 < l2 and / or l3 < l4, and the difference between the required thrust vectors F1 and F2 or F3 and F4 of the propulsion devices 1A1, 1A2 or 1B3, 1B4 further increases.

[0214] (ii) In the design of the second case, the two propulsion devices 1A1 and 1A2 are preferably designed to be the same, particularly from a structural point of view, and / or the two propulsion devices 1B3 and 1B4 are designed to be the same from a structural point of view. That is, they are structurally the same. For example, they have the same size, the same span, the same number of rotor blades, and / or generate similar or the same (maximum) propulsion force / thrust vector. In this case, therefore, F1 = F2 (or F1 ≈ F2) and / or F3 = F4 (or F3 ≈ F4).

[0215] From equation (25a) and equation (25b),

Number

[0216] With respect to the forward flight direction, the center of gravity S, 250 of the aircraft is determined from the (geometric) center l / 2 between the opposing propulsion devices 1A1, 1A2 and 1B3, 1B4, respectively, along the first direction 201 and / or the second direction according to equations (24a) and (24b), respectively:

number

[0217] Now, if the aircraft is configured with structurally identical and equal-sized propulsion devices 1A1, 1A2 and / or propulsion devices 1B3, 1B4 for each pair of propulsion devices 1A1, 1A2 or 1B3, 1B4, and therefore with thrust / thrust vectors of approximately equal magnitude, the center of gravity S, 250 can therefore be optimally positioned so that the torques generated by the propulsion devices are purely balanced by the position of the center of gravity S, 250. Said optimal position is determined by equation (27a) and / or equation (27b).

[0218] It should be noted here and below that for equally rotating propulsion devices 1B3, 1B4, only the position of the center of gravity in the first direction 201 plays a role in the considerations. The support or positioning of the center of gravity relative to the second direction and the vertical direction 203 is irrelevant here and is left to the discretion of those skilled in the art. Therefore, for equally rotating propulsion devices 1A1, 1A2, only the position of the center of gravity in the second direction plays a role in the considerations. In this case, the support or positioning of the center of gravity relative to the first direction 201 and the vertical direction 203 is irrelevant. However, if the aircraft utilizes the positive effect of the Magnus effect both when moving forward in the first direction and when moving forward in the second direction, only its position relative to the vertical direction 203 remains freely selectable, since the optimal position of the center of gravity is determined by both equation (27a) and equation (27b).

[0219] (iii) According to the present invention, it is also possible to combine the aspects of the design (i) of the first case with the aspects of the design (ii) of the second case, i.e., the center of gravity S, 250 of the aircraft can be shifted from the geometric center of the aircraft fuselage 220 in such a way that the conditions (17), (20), (25a), (25b) for stable hover flight at certain predetermined and even different thrust vectors / thrusts of the individual propulsion devices are met.

[0220] In practical applications, it is not always possible to arrange the mass of the aircraft so that the overall center of gravity S,250 can be positioned exactly at the predetermined optimum position described in (i), (ii), or (iii) (in case design (i), l = l and / or l = l; for case design (ii), see equations (26a), (26b), (27a), and (27b)). Therefore, we define below the range within which the center of gravity S,250 can be positioned so that it is still possible to support torque compensation by the pair of thrust vectors F,2001, F,2002 of propulsion devices 1A1, 1A2 and torque compensation by the pair of thrust vectors F,2003, F,2004 of propulsion devices 1B3, 1B4, respectively.

[0221] For this purpose, first, one of the propulsion devices 1A1, 1A2, 1B3, 1B4 is driven by a maximum allowable (usually predetermined) thrust force / maximum allowable thrust vector F i,max Suppose we can generate F i,max is the optimal configuration F i,opt (as also already explained in more detail in relation to the first aspect of the invention).

[0222] Considering the momentum theorem according to equation (17), first, F 1,opt ≦F 1,max ⇒ F 2,opt ≧F 2,min ≡F s -F 1,max -F 34,opt is obtained, Therefore, the maximum allowable thrust vector ratio

number

[0223] F 2,opt ≦F 2,max ⇒ F 1,opt ≧F 1,min ≡F s -F 2,max -F 34,opt in the case of, This is the minimum allowable thrust vector ratio

number

[0224] In equations (17) and (20), the boundary condition in equation (23), F 12 =F 34 Using

number

[0225] These thrust vector ratios F1 / F2 are also described by equation (25a), and using equation (25a), the maximum allowable distance in the first direction 201 from the geometric center of the forward propulsion device 1A to the center of gravity S, 250 in forward flight 1 is

number

number

[0226] The center of gravity S, 250 in the first direction 201 is

number

[0227] Equation (24a) also allows specifying the tolerance range (28) in the first direction relative to the axis of rotation of the propulsion devices 1B3, 1B4 or a line passing through the propulsion devices 1B3, 1B4. The range specification is therefore based on the distance l 34 and the corresponding limit l 34,min and limit l 34,max This is done with the help of

[0228] Similarly, the tolerance of the center of gravity S, 250 in the second direction (here, the y direction) is

number

number

number

[0229] Equation (24b) also allows specifying the tolerance range (29) in a second direction relative to the axis of rotation of the propulsion devices 1A1, 1A2 or a line passing through the propulsion devices 1A1, 1A2. The range specification is therefore based on the distance l 12 and the corresponding limit l 12,min and limit l12,max This is done with the help of

[0230] Figure 7d serves to explain the above range in which the centre of gravity S, 250 of the aircraft can be suitably positioned to implement the invention according to the second aspect. Figure 7d shows a propulsion device 1A corresponding to that described in relation to Figures 7a and 7b. 1、 1A2 and 1B 3、 1B4. 1、 1A 2、 1B 3、 1B4 are designed identically in terms of structure (see case design (ii) above), and here, in particular, a1 = a2 = a3 = a4 = a 12 =a 34 Assume that ≡a.

[0231] First, we further assume that torque compensation is realized purely by the location of the center of gravity, S,250,. F 1,opt =F 2,opt =F opt ⇒ M 1,opt =M 2,opt =M opt applies. In the example aircraft considered here, the total weight force generated by the corresponding total mass is: F s =1000N and the ratio / proportionality coefficient is typically a=0.2m, The distance of the propulsion device in the first direction (x direction in Figures 7a, 7b) is l=l1+l2=2m It is defined as follows.

[0232] Based on these specifications, the optimal center of gravity position is calculated using equations (25a) and (26a).

number

[0233] The aircraft's overall center of gravity S,250 is located at l 1,opt = 1.1 m, we now define the range within which the position of the center of gravity S,250 can be positioned so that torque compensation can be achieved by the thrust / thrust vectors of the propulsion devices 1A1, 1A2, 1B3, 1B4. For this purpose, for convenience, the maximum allowable thrust that can be generated by each of the propulsion devices 1A1, 1A2, which is positioned in the first direction and controlled by the thrust vector control, is F 1,max =F 2,max =275N It is defined as follows.

[0234] With this specification, and the boundary condition according to equation (23),

number

number

number

[0235] Assuming the aircraft is designed symmetrically, the same value for the tolerance of l3 is obtained. When both conditions are considered, the center of gravity S,250 is properly positioned within a square area determined by the specified limits with respect to the plane defined by the propulsion devices and the aircraft fuselage. There are no restrictions on vertical positioning.

[0236] Finally, it is further noted that the second aspect of the invention is not limited to aircraft with four propulsion devices: for example, more than two propulsion devices may be arranged along one direction, or several propulsion devices may be arranged in lines parallel to one another.

[0237] We now generalize equations (17), (18), (20), and (21) for an aircraft according to the invention having n propulsion devices 1C (n>2). Figure 8a shows a plan view of such an aircraft, and Figure 8b shows a side view of the aircraft. We assume that the mathematical and physical description of the aircraft is carried out 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 of z=0. The propulsion devices 1C are arranged around the aircraft fuselage 220 in the plane of z=0 (star-shaped). The origin O of the coordinate system is placed at the geometric center of the aircraft. Then, r i Let ,i∈{1,…,n} be the position vector for the i-th thrust vector of the corresponding propulsion device 1C. Let s be the position vector for the aircraft's center of gravity S,250. The aircraft's weight force vector is F s =(0,0,F s ) for the stable hovering flight considered here, the thrust vector is F i ,i∈{1,…,n}, F i ,=(0,0,-F i ),i∈{1,…,n} is.

[0238] Finally, in hover flight, the propulsion device operates at an angular velocity ω i,i∈{1,…,n} and we are dealing with vectors lying in the xy plane. The torque that must be applied by the aircraft, already detailed in the introduction, is therefore given by the relation M i =a i *F i It can be calculated by taking into account

number

[0239] The equilibrium conditions for equations (17), (18), (20), and (21) are therefore:

number

[0240] From the angular momentum balance theorem, the position vector s of the center of mass S,250 can be determined as follows:

number

number

number

number

[0241] Equation (32) provides two conditions, one on the x-components of the torques and the other on their y-components. Equations (31) and (32) (taking into account the relationship in equation (30)) therefore provide the magnitudes F of the n thrust vectors. i and three conditions on the two coordinates of the center of gravity. This leaves n+2-3=n-1 n thrust vectors that can be specified. Therefore, in the more general case considered here, the appropriate range for the plane in which the propulsion devices 1C and the aircraft fuselage 220 lie can also be determined by varying the thrust vector of one or more of the n propulsion devices 1C and requiring the center of gravity S, 250 to be positioned (see equation (33)) so that the torque compensation according to equation (32) can be compensated by the thrust / thrust vector of the propulsion device. For this purpose, it may be useful to drive one or more of the propulsion devices at the maximum allowable thrust.

[0242] It is advantageous to consider the first direction and / or the second direction (along which the propulsion devices rotate in substantially the same direction) as directions perpendicular to two predetermined forward flight directions, in that the optimal position of the center of gravity of the considered configuration is determined by the intersection of two straight lines, in which case the center of gravity is therefore preferably displaced from the geometric center in (i) a direction lying in a plane defined by the propulsion devices and the aircraft fuselage that is perpendicular to the first direction, and / or in (ii) a direction lying in a plane defined by the propulsion devices and the aircraft fuselage that is perpendicular to the second direction.

[0243] 9a shows an embodiment according to a second aspect of the invention in which three propulsion devices 1C1, 1C2, 1C3 are arranged around the aircraft fuselage 220 of an aircraft to form the sides of an equilateral triangle. Propulsion devices 1C1 and 1C2 are shown arranged on a line g1, which thereby defines a first direction according to the invention. In the illustrated embodiment, propulsion device 1C3 is arranged on a line g2 that is perpendicular to line g1 and passes through the geometric center G of the aircraft, in this case the geometric center G of the equilateral triangle. Line g2 defines a second direction according to the invention. Here, the axes of rotation 5C1, 5C2 of propulsion devices 1C1, 1C2, and 1C3 are aligned along the line g2. 2、 5C3 each point towards (or away from) the geometric center G. In the illustrated embodiment, only the rotation axis 5C3 is aligned exactly parallel to the second direction defined by g2. The rotation axes 5C1, 5C2 are not exactly parallel to the first direction defined by g1. As can be seen with the help of simple geometric considerations, the rotation axis 5C1 includes an angle α1 = 30° with the line g1 (first direction), and similarly, the rotation axis 5C2 includes an angle α2 = 30° with the line g1 (first direction). Such angles are included within the concept of rotation axes substantially aligned with the first direction according to the present invention. However, the angles can also preferably be selected to be smaller. Propulsion device 1C 1、 If the propulsion devices 1C1, 1C3 rotate about their associated rotation axes 5C1 and 5C3 in substantially the same direction of rotation as defined above, then the success of the present invention also occurs when the aircraft moves in particular along the second direction defined by g2. If the propulsion devices 1C1, 1C3 rotate about their associated rotation axes 5C1 and 5C3 in substantially the same direction of rotation, then the advantages of the present invention have a positive effect in particular when the aircraft moves along the angle bisector 1C1-G-1C3.

[0244] 9b shows an aircraft according to a second aspect of the invention in which seven propulsion devices 1C1,...,1C7 are arranged in one plane around the aircraft fuselage 220. The propulsion devices 1C1,...,1C7 are arranged to form the vertices of a regular heptagon. Each propulsion device is rotatably mounted about an associated axis of rotation 5C1,...,5C7. In the illustrated embodiment, the axes of rotation 5C1,...,5C7 point towards the aircraft and the geometric centre G of the heptagon, respectively. This embodiment uses an (odd) n=2j+1 (j>1) number of propulsion devices 1C1,...,1C 2j+1 are arranged around the aircraft fuselage 220 so as to form the vertices of a regular polygon with (2j+1) sides. 2j+1 should point towards (or away from) the geometric center G. In this case, the first straight line g1 is drawn by the two propulsion devices 1C1 and 1C2. (n+1) / 2 This straight line g1 defines a first direction according to the invention. k and 1C k+(n-1) / 2 ,

number

number

[0245] With the help of simple geometric considerations, each of the rotation axes 5C1 and 5C (n+1) / 2 is the angle α1 = α (n+1) / 2 =90° / n, rotation axis 5C k and rotation axis 5C k+(n-1) / 2 and the angle α between the line g2 k =α k+(n-1) / 2= 90° / n. For the heptagon shown, therefore, we have α1 = α3 = α4 = α6 = 90° / 7 ≈ 12.86°. For a regular polygon with (2j+1) sides, it is therefore advantageous if the rotation axes of the propulsion devices, which lie on the lines g1 and g2 defining the first and second directions, respectively, subtend angles between 0° and 90° / n with the associated lines g1 and g2, respectively.

[0246] The angle ε between g1 and g2 can be easily calculated using geometric relations.

number

[0247] 9c shows an aircraft according to a second aspect of the invention, in which six propulsion devices 1C1,...,1C6 are arranged in a plane around the aircraft fuselage 220. The propulsion devices 1C1,...,1C6 are arranged to form the vertices of a regular hexagon. Each propulsion device is rotatably mounted about an associated axis of rotation 5C1,...,5C6. In the illustrated embodiment, the axes of rotation 5C1,...,5C6 point towards the aircraft and the geometric centre G of the hexagon, respectively. This example shows an even number of n=2j (j>1) propulsion devices 1C1,...,1C6. 2j , 5C1, ..., 5C2 are arranged around the periphery of the aircraft fuselage 220 so that they form the vertices of a regular polygon with 2j sides. The aircraft fuselage 220 is located between each pair of opposing propulsion devices of the regular polygon with 2j sides. The associated axes of rotation 5C1, ..., 5C3 2j are oriented towards (or away from) the geometric center G. In this case, the two propulsion devices 1C1 and 1C n / 2+1 It is convenient to consider a first straight line g1 as passing through the two propulsion devices 1C, which defines a first direction according to the invention. k and 1C k+n / 2 ,

number

[0248] In the illustrated embodiment, the rotation axes of the propulsion devices located on the lines g1 and g2 are aligned (mathematically precisely) parallel to the first direction and the second direction, respectively.

[0249] Particularly preferably, the first direction and the second direction are substantially perpendicular to each other, in particular perpendicular, which is always possible if the propulsion device forms the vertices of a polygon with 4j sides.

[0250] The angle ε between g1 and g2 (i.e., the first direction and the second direction) for the above 2j vertices can be easily deduced by using the geometric relationship:

number

[0251] From the above embodiments, it can be seen that for the arrangement of any number (even or odd) of propulsion devices at the vertices of a regular polygon with n sides, it is sufficient if the rotation axes of the propulsion devices, which lie on the lines g1 and g2 defining the first and second directions, respectively, form angles with the corresponding lines g1 and g2, which angles are between 0° and 30° (if n>2), particularly preferably between 0° and 18° (if n>3); furthermore, it is advantageous if the lines g1 and g2 (and thus the first and second directions) are selected so that the angle between them is greater than or equal to 60°, in particular in the range of 60° to 90°.

[0252] Appendix (Derivation of the relationship between thrust and power) The derivation of thrust and power is based on momentum theory, and the propulsion device / rotor is considered to be an actuator disk without information about the number and shape of the rotor blades. The flow is simplified by defining it as one-dimensional, unsteady, incompressible, and frictionless, resulting in the corresponding laws of conservation of mass, momentum, and energy. In what follows, all quantities within the actuator disk plane are given the additional subscript a, all quantities far above the actuator disk plane (inflow plane) have the additional subscript 0, and all quantities far below the actuator disk plane (outflow plane) are given the additional subscript ∞.

[0253] Law of conservation of mass: Due to the flow assumptions, mass flow follows the law of conservation of mass.

number

[0254] Conservation of momentum: Due to the flow assumptions, the driving force follows the law of conservation of momentum.

number

[0255] Since the rotor does not affect the inflow plane, v i0 =0, and from here,

number

[0256] Law of Conservation of Energy: Flow assumptions and v i0= 0, the power or work done per unit time relative to the actuator disk plane follows the law of conservation of energy.

number

[0257] By inserting the mass flow into the actuator disc plane,

number

[0258] With propulsion, the power is

number

[0259] Directly from these equations, the relationship

number

[0260] List of Reference Numbers 100 Aircraft according to a first aspect of the invention 120 Aircraft body Propulsion device located in the front area of ​​1F 1R Propulsion device located in the rear area 101 longitudinal direction of aircraft 100 102 transverse direction of aircraft 100 103 Vertical direction of aircraft 100 121 Aircraft 100 Bow / Nose 122 Tail of Aircraft 100 1 Propulsion Device 2. Propulsion device rotor blades 3 Pitch mechanism 31 Coupling Device 32 Connection point 33 Bearing Device 4 Offset Device 11 Propulsion Device 1 Disk 5 Rotation axis of the propulsion device 51 Direction of rotation of propulsion device 52 Radius of propulsion device 61 Connecting Elements 7, 71 Force / thrust vector acting on the propulsion device 72 Magnus effect contribution to thrust vector 8 Torque on the propulsion device 9 Air Inflow 110 Arrow indicating the direction of aircraft movement 150 Center of gravity of aircraft 100 701 Total thrust vector generated by propulsion device 1F 702 Total thrust vector generated by propulsion device 1R 81 Total torque generated by propulsion device 1F 82 Total torque generated by propulsion device 1R 131 the distance in the longitudinal direction between the center of gravity 150 and the propulsion device 1F 132 the longitudinal distance between the center of gravity 150 and the propulsion device 1R 160 aircraft weight g i The i-th line where the propulsion device is located n i straight g i Number of propulsion devices placed along N total number of lines K Total number of propulsion devices F ij straight g i Thrust vector generated by the j-th propulsion device placed on F i straight gi Thrust vector generated by all propulsion devices placed on M i straight g i Torque generated by all propulsion devices placed above x i line g i Longitudinal coordinate of X S Longitudinal coordinate of center of gravity 150 200 Aircraft according to a second aspect of the present invention 220 aircraft fuselage 1A, 1B, 1C, 1A1, 1A2, 1B3, 1B4, 1C i Aircraft Propulsion Device 200 221, 222 Arms for connecting the propulsion devices 1A, 1B to the aircraft fuselage 220 201 First Direction 202 Second Direction 203 Vertical 5A Rotation axis of propulsion device 1A 5B Rotation axis of propulsion device 1B 5C i Propulsion Device 1C i Rotation axis of α i Rotation axis 5C i and the angle between the first direction or the second direction ε is the angle between the first direction and the second direction 250 Center of gravity of aircraft 200 G geometric center O Origin of the coordinate system 2001, 2002 Thrust vectors generated by the propulsion device 1A1 and the propulsion device 1A2, respectively 2003, 2004 Thrust vectors generated by thrust device 1B3 and thrust device 1B4, respectively 2012 Total thrust vector generated by propulsion devices 1A1 and 1A2 2034 Total thrust vector generated by propulsion devices 1B3 and 1B4 230 Thrust vector / distance between geometric centers of propulsion devices 1A1 and 1A2 231 Distance from aircraft center of gravity 250 to thrust vector 2001 232 Distance from aircraft center of gravity 250 to thrust vector 2002 234 Center of gravity 250 and thrust vector F 34 , 2034 / the rotation axis of the propulsion device 1B3, 1B4 / the distance between the straight line passing through the propulsion device 1B3, 1B4 235 Thrust vector / distance between geometric centers of propulsion devices 1B3 and 1B4 236 Distance from aircraft center of gravity 250 to thrust vector 2003 237 Distance from aircraft center of gravity 250 to thrust vector 2004 239 Center of gravity 250 and thrust vector F 12 ,2012 / the distance between the rotation axis of the propulsion device 1A1, 1A2 / the straight line passing through the propulsion device 1A1, 1A2 251 Direction of rotation of propulsion devices 1B3, 1B4 280 Total torque generated by propulsion devices 1B3, 1B4 285 Total torque generated by propulsion devices 1A1 and 1A2

Claims

1. An aircraft (100), an aircraft body (120) defining a longitudinal direction (101), a vertical direction (103) and a transverse direction (102), said longitudinal direction corresponding to the direction from the tail (122) to the nose (121) of said aircraft (100), said vertical direction corresponding to the direction of the Earth's gravitational force when said aircraft (100) is resting on the ground, and said transverse direction being perpendicular to said longitudinal direction and said vertical direction; at least two propulsion devices (1, 1F, 1R) each rotatable about an associated axis of rotation (5) and generating a respective associated thrust vector (701, 702), each consisting of a cyclogyro rotor; a first number of the propulsion devices (1F) are arranged along a first straight line parallel to the transverse direction (102), and a second number of the propulsion devices (1R) are arranged along a second straight line parallel to the transverse direction (102); further, the first number of the propulsion devices (1F) are arranged in a front region of the aircraft with respect to the longitudinal direction (101), and the second number of the propulsion devices (1R) are arranged in a rear region of the aircraft with respect to the longitudinal direction (101); the first straight line is spaced apart from the second straight line; the center of gravity of the aircraft (100) is located between the first line and the second line with respect to the longitudinal direction (101); at least two propulsion devices (1, 1F, 1R); Including, 1. The aircraft (100) adapted to perform hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially vanish with respect to the center of gravity (150) of the aircraft, In the hover flight, each of said associated axes of rotation (5) is oriented substantially in said transverse direction (102) of the aircraft body; - each of said at least two propulsion devices (1, 1F, 1R) rotates in substantially the same direction of rotation (51) around said respective associated axis of rotation (5); characterized in that Aircraft (100).

2. The center of gravity (150) of the aircraft when performing hover flight is at a longitudinal distance l from the line on which the propulsion devices (1F) of the front region are arranged. １ placed in [Equation 1] and [Equation 2] and During the ceremony, R ｍｉｎ is the minimum allowable ratio between the thrust vector (701) of the propulsion device (1F) located in the one of the forward regions and the thrust vector (702) of the propulsion device (1R) located in the other of the aft regions, R ｍａｘ is the maximum allowable ratio between the thrust vector (701) of the propulsion device (1F) located in the one of the forward regions and the thrust vector (702) of the propulsion device (1R) located in the other of the aft regions, l is the distance between the first line and the second line, a １ is the characteristic number of the propulsion devices (1F) located in the front region, a ２ is the characteristic number of the propulsion device (1R) located in the rear region, The aircraft (100) of claim 1.

3. 3. The aircraft (100) of claim 1 or 2, adapted to have the associated axes of rotation aligned parallel during hover flight.

4. An aircraft (200), an aircraft fuselage (220), at least three propulsion devices (1A, 1B) mounted substantially planarly around the aircraft fuselage, rotatable about respective associated axes of rotation (5A, 5B), generating respective associated thrust vectors (2001, 2002, 2003, 2004), and consisting of cyclogyro rotors; Including, The aircraft (200) is adapted to perform hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially disappear with respect to a center of gravity (250) of the aircraft, and during the hover flight: the associated rotation axes (5A) of two of the at least three propulsion devices (1A, 1B) are substantially aligned in a first direction (201) and the associated rotation axes (5B) of the other of the at least three propulsion devices (1A, 1B) are substantially aligned in a second direction (202); - said first direction (201) is not parallel to said second direction (202); - said first direction (201) and said second direction (202) are in the same plane; - each of the two propulsion devices (1A) with their rotation axes (5A) oriented in the first direction (201) rotates in substantially the same rotational direction around their associated rotation axes (5A) during hover flight; An aircraft (200).

5. at least four propulsion devices (1A, 1B) mounted around the periphery of said aircraft fuselage, rotatable about respective associated axes of rotation (5A, 5B), and generating respective associated thrust vectors (2001, 2002, 2003, 2004); Including, The aircraft (200) is adapted to perform the hover flight, and during the hover flight: the associated rotation axes (5A) of two of the at least four propulsion devices (1A, 1B) are substantially aligned in the first direction (201) and the associated rotation axes (5B) of a further two of the at least four propulsion devices (1A, 1B) are substantially aligned in the second direction (202); - during hover flight, each of the two propulsion devices (1A) whose rotation axes (5A) are oriented in the first direction (201) rotate in substantially the same rotational direction about their associated rotation axes (5A), and / or during hover flight, each of the two propulsion devices (1B) whose rotation axes (5B) are oriented in the second direction (202) rotate in substantially the same rotational direction about their associated rotation axes (5B), The aircraft (200) of claim 4.

6. three propulsion devices arranged around the periphery of the aircraft fuselage to form sides of an equilateral triangle; - said aircraft fuselage (220) is located at the geometric center of said equilateral triangle; - said first direction is defined by a straight line along which two of said three propulsion devices are located; - said second direction is substantially perpendicular to said first direction; the rotation axis of each of the two propulsion devices located on the straight line pointing in the first direction includes an angle with the straight line ranging from 0° to 30°; The aircraft (200) of claim 4.

7. n propulsion devices are arranged around the periphery of the aircraft fuselage so as to form vertices of a regular polygon having n sides, where n>3; - said aircraft fuselage (220) is located at the geometric center of said regular polygon having n sides; - said first direction is defined by a first straight line along which two of said n propulsion devices are located; - the second direction is defined by a second straight line along which two further of the n propulsion devices are located; the rotation axis of each of the two propulsion devices located on the first straight line pointing in the first direction includes an angle with the first straight line in the range of 0° to 18°; The aircraft (200) of claim 5.

8. The aircraft (200) of claim 7, wherein the angle between the first line and the second line ranges from 72° to 90°.

9. - said second direction (202) is substantially perpendicular to said first direction (201); two of the at least four propulsion devices (1A) are arranged along the first direction (201) and two of the at least four propulsion devices (1B) are arranged along the second direction (202) substantially perpendicular to the first direction; An aircraft (200) according to any one of claims 5, 7 and 8.

10. The center of gravity (250) of the aircraft when performing hover flight is a distance l in the first direction (201) from a line on which the propulsion device (1B) is arranged in the second direction (202). ３４ placed in [Equation 3] and [Equation 4] and During the ceremony, R ｍｉｎ is the minimum allowable ratio between the thrust vectors (2001, 2002) of the propulsion devices (1A) arranged along the first direction, R ｍａｘ is the maximum allowable ratio between the thrust vectors (2001, 2002) of the propulsion devices (1A) arranged along the first direction, a ３４ is the characteristic number of the propulsion devices (1B) arranged along the second direction (202), l is the distance between the geometric centers of the propulsion devices (1A) arranged in the first direction, and / or The center of gravity (250) of the aircraft when performing hover flight is a distance l in the second direction (202) from a line along which the propulsion device (1A) is arranged in the first direction (201). １２ placed in [Equation 5] and [Equation 6] and [Equation 7] is the minimum allowable ratio between the thrust vectors (2003, 2004) of the propulsion devices (1B) arranged along the second direction (202), [Equation 8] is the maximum allowable ratio between the thrust vectors (2003, 2004) of the propulsion devices (1B) arranged along the second direction (202), a １２ is the characteristic number of the propulsion devices (1A) arranged along the first direction (201), l' is the distance of the geometric center of the propulsion device (1B) arranged in the second direction; The aircraft (200) of claim 9.

11. 11. The aircraft (100, 200) of any one of claims 1 to 10, further adapted such that, during hover flight, when one or more of the propulsion devices generate a respective particular predetermined thrust vector associated with one or more of the propulsion devices, the center of gravity (150, 250) of the aircraft is positioned relative to the center of gravity of the aircraft to substantially eliminate all forces acting on the aircraft and all torques acting on the aircraft.

12. The aircraft (100, 200) of any one of claims 1 to 11, wherein the aircraft is further adapted to perform hover flight with a particular predetermined thrust vector associated therewith that is approximately equal.

13. The aircraft (100, 200) of any one of claims 1 to 12, further comprising a displacement device for displacing the center of gravity (150, 250) of the aircraft.

14. further comprising a fuel tank for supplying fuel to the propulsion device and / or a battery for supplying power to the propulsion device; 14. The aircraft (100, 200) of claim 13, when relying on claim 11 or claim 12, wherein the displacement device is adapted to reposition fuel from the fuel tanks or the batteries within the aircraft, thereby positioning the center of gravity (150, 250) such that the aircraft performs hover flight when one or more of the propulsion devices generate the respective specific predetermined thrust vectors.

15. The aircraft (100, 200) of any one of the preceding claims, comprising thrust vector controls for individually controlling the thrust vectors of the propulsion devices.

16. A method of manufacturing an aircraft (100) according to any one of claims 1 to 3 or 11 to 15, comprising the following steps: - each of said associated axes of rotation (5) is oriented substantially in said transverse direction (102) of said aircraft body (120); - when each of said at least two propulsion devices (1F, 1R) rotates in substantially the same direction of rotation (51) around said respective associated axis of rotation (5), positioning the center of gravity (150) of the aircraft such that one or more of the propulsion devices (1F, 1R) generate specific predetermined thrust vectors associated with one or more of the propulsion devices (1F, 1R), respectively, to cause the aircraft to perform a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially vanish with respect to the center of gravity of the aircraft; A method comprising:

17. A method of manufacturing an aircraft (200) according to any one of claims 4 to 15, comprising the following steps: the associated rotation axes (5A) of two of the at least three propulsion devices (1A, 1B) are substantially aligned in the first direction (201) and the associated rotation axes (5B) of the other of the at least three propulsion devices (1A, 1B) are substantially aligned in the second direction (202); - when each of the two propulsion devices (1A) with their axes of rotation oriented in the first direction (201) rotates in substantially the same rotational direction about their associated axes of rotation (5A) during hover flight, positioning the center of gravity (250) of the aircraft such that one or more of the propulsion devices (1A, 1B) generate specific predetermined thrust vectors associated with one or more of the propulsion devices (1A, 1B), respectively, to cause the aircraft to perform a hover flight in which all forces acting on the aircraft and all torques acting on the aircraft substantially vanish with respect to the center of gravity (250) of the aircraft; A method comprising:

18. A method of controlling an aircraft (100), comprising: an aircraft body (120) defining a longitudinal direction, a vertical direction and a transverse direction, the longitudinal direction corresponding to the direction from the tail (122) to the nose (121) of the aircraft, the vertical direction corresponding to the direction of the Earth's gravitational force when the aircraft (100) is resting on the ground, and the transverse direction being perpendicular to the longitudinal direction and the vertical direction; at least two propulsion devices (1, 1F, 1R) rotatable about respective associated axes of rotation (5) and generating respective associated thrust vectors, at least two propulsion devices (1, 1F, 1R) consisting of cyclogyro rotors, a first number of the propulsion devices (1F) are arranged along a first straight line parallel to the transverse direction (102), and a second number of the propulsion devices (1R) are arranged along a second straight line parallel to the transverse direction (102); further, the first number of the propulsion devices (1F) are arranged in a front region of the aircraft with respect to the longitudinal direction (101), and the second number of the propulsion devices (1R) are arranged in a rear region of the aircraft with respect to the longitudinal direction (101); the first straight line is spaced apart from the second straight line; the center of gravity of the aircraft (100) is located between the first line and the second line with respect to the longitudinal direction (101); at least two propulsion devices (1, 1F, 1R); and The following steps: determining thrust vectors associated with the at least two propulsion devices such that the aircraft performs hover flight when each of the rotational axes associated with the at least two propulsion devices is oriented substantially in the transverse direction of the aircraft body and when each of the at least two propulsion devices rotates in substantially the same rotational sense about its associated rotational axis, During hover flight, all forces acting on the aircraft and all torques acting on the aircraft relative to the center of gravity (150) of the aircraft substantially disappear. determining an associated thrust vector; - driving each of said propulsion devices (1F, 1R) in substantially the same sense of rotation so that each of said propulsion devices generates a specific associated thrust vector; Including, method.

19. A method of controlling an aircraft (200), comprising: an aircraft fuselage (220), at least three propulsion devices (1A, 1B) mounted substantially planarly around the aircraft fuselage and rotatable about respective associated axes of rotation (5A, 5B) and generating respective associated thrust vectors; and The following steps: determining the associated thrust vectors such that the aircraft performs a hover flight when two of the rotation axes (5A) associated with the at least three propulsion devices (1A, 1B) are substantially aligned in a first direction (201) and rotate in substantially the same rotational sense about their respective associated rotation axes (5A), and / or when the other of the rotation axes (5B) associated with the at least three propulsion devices (1A, 1B) are substantially aligned in a second direction (202) that is not parallel to the first direction, During hover flight, all forces acting on the aircraft and all torques acting on the aircraft substantially disappear relative to the aircraft's center of gravity (250). determining an associated thrust vector; - aligning the associated rotation axes (5A) of two of the at least three propulsion devices (1A, 1B) substantially in the first direction (201) and the associated rotation axis (5B) of another of the at least three propulsion devices (1A, 1B) substantially in the second direction (202), where the first direction (201) and the second direction (202) are in the same plane; - driving each of said propulsion devices (1A, 1B) so that it rotates in an associated rotational direction and generates a particular associated thrust vector, each of the two propulsion devices (1A) having a rotation axis substantially oriented in the first direction rotates in substantially the same rotational direction about its associated rotation axis (5A); a driving step; A method comprising:

20. 20. A method of controlling an aircraft (100, 200) according to claim 18 or 19, wherein all of the determined associated thrust vectors are selected to be substantially identical.

21. positioning the center of gravity (150, 250) of the aircraft such that, when the propulsion device generates a particular predetermined thrust vector associated with the propulsion device, all forces acting on the aircraft and all torques acting on the aircraft are substantially nullified relative to the center of gravity of the aircraft; The method of controlling an aircraft (100, 200) according to any one of claims 18 to 20, further comprising:

22. An aircraft (100, 200) according to any one of the preceding claims, wherein each of said propulsion devices (1, 1F, 1R, 1A, 1B) is structurally identical.

Images