DE102022114599A1 - aircraft - Google Patents

Abstract

For mission profiles of aircraft that are hovering for most of the mission, according to the current state of technology, airships that require a lot of space or energy-intensive aircraft with multiple rotors are mainly used. This aircraft only has a single rotor assembly with one or several rotors that can be mechanically firmly coupled together, with the rotors rotating about the transverse axis, such as cyclorotors. The aircraft is stabilized around the transverse axis by a counter-torque (pitching moment) which is generated by the fuselage and the payload to be transported. To avoid instabilities, the overall center of gravity of the fuselage and payload is below and at a distance from the rotor. Due to the fact that only one motor is required for propulsion, the rotors are arranged concentrically to a single rotor shaft, there is no need to have a gap between the wing tips of two rotors, and no cabin or structure needs to interact with the wake, greater energy efficiency is imminent Comparison to aircraft based on the current state of the art is to be expected. A particularly high level of controllability is to be expected for a rotor assembly with two or more rotors. The invention relates to a rotary-wing aircraft for transporting a load, in particular in the form of a person or cargo, as well as the use as a crane helicopter and aviation sports equipment.

Description

The invention generally relates to rotary wing aircraft that are capable of hovering flight.

STATE OF THE ART

It is known that fixed-wing thrust vector control aircraft (VTOL aircraft), ornithopters, multicopters and helicopters can achieve and maintain the hover flight condition.

Because of their small rotors or turbines, VTOL aircraft are relatively energy inefficient when hovering. This is mainly the case because the design of VTOL aircraft assumes that they only need to achieve short hover periods before and after longer cross-country flights. The most well-known and most energy-efficient VTOL aircraft when hovering are the tiltrotor and tilt-wing aircraft.

Ornithopters are nature-inspired flying machines developed by humans, although these flying machines are also energy inefficient. In particular, the motors that alternately change their direction of rotation are energy inefficient. In addition, ornithopters that can maintain hovering flight can only be implemented in small, light models (https://www.dailymail.co. uk/sciencetech/article-6165071 /Scientists-develop-newflying-robot-mimics-flight-fruit-flies .html and https://www.youtube.com/watch?v=dURvisvMYT8), which does not qualify these aircraft for practical use. This is mainly because the principle of the ornithopter depends on a flow around it at low Reynolds numbers.

The fact that multicopters have multiple rotors makes these aircraft inefficient not only from an aerodynamic perspective, but also from a drive perspective. The aerodynamic inefficiency is mainly due to suboptimal use of the rotor area and rotor-rotor flow interactions, including in the wake. The inefficiency of the propulsion can be attributed to the lower electrical energy efficiency of the smaller engines compared to a rotorcraft with only one rotor (and therefore a single larger engine). In addition, multicopters usually do not have a main control mechanism as is the case with conventional rotorcraft (collective/cyclic control), which is why they have to adjust the individual rotation speeds of individual motors in order to keep the aircraft stable in the air, which is energy inefficient. Cyclorotor multicopters are an alternative. These cyclorotor multicopters suffer less from the lack of controllability of the individual rotors, but they still have similar energy efficiency problems as other multicopters.

The conventional helicopter is the most energy efficient for the hover flight condition. The helicopter is also the preferred aircraft for carrying heavy loads while hovering. However, the concept suffers from high control complexity and power losses due to an unavoidable drive (tail rotor, exhaust nozzles, coaxial rotor or similar) for torque compensation, as well as the constant flow interaction between the wake and the cabin or payload.

The following paragraphs will describe the cyclorotor in more detail in terms of its advantages, while also better defining the limitations of the helicopter. The aim is to show that cyclorotors are more energy efficient and easier to control at certain flight levels, especially when hovering.

In general, a cyclorotor, cyclocopter, cyclologyro, cycloidal propeller or cycloidal rotor describes a cylindrical rotor whose cylinder jacket consists of several rotor blades, which are oriented with their span parallel to the rotor axis and rotate about this axis on the cylinder jacket. In the historical patents US 1753252 A and US 2045233 A The general operating principle of a cyclorotor is described.

A blade pitching mechanism enables the generation and control of a thrust vector by periodically changing the angle of attack of the rotor blades. The rotor blades are pivoted along their span and their angle of attack is varied individually using control rods.

The advantage of the cyclorotor is that the rotational movement of individual rotor blades of a cyclorotor around the longitudinal axis of the blade (running from the blade root to the blade tip) encounters significantly lower resistance due to the mass moment of inertia than in a conventional helicopter For a comparable change of direction maneuver, at least the entire rotor and often also the entire cabin must be moved.

Since only the blade angle of attack has to be varied to change the direction of flight of a cyclorotor, shorter control times can be achieved with the same energy expenditure due to the lower mass moment of inertia, which allows the aircraft to gain considerable agility. It should be noted that the cyclorotor aircraft does not necessarily have to have a lower mass moment of inertia in terms of its overall volume compared to a conventional helicopter, but simply benefits from the fact that less mass has to be moved in order to achieve a change in thrust vector.

Due to the arrangement of the rotor blades of a cyclorotor, the speed distribution (and therefore the Mach number) is constant along the blade span, allowing operation with lower losses compared to conventional rotary wing aircraft.

Furthermore, cyclorotors operate at lower rotational speeds compared to conventional helicopters, which leads to a significant reduction in noise. The property of low noise characteristics as an advantage of cyclorotors has already been established and published at SJ Kim: “Design, development and flight test of vertical take-off and landing UAV, cyclocopter”, Guggenheim Memorial Lecture, 30th ICAS Congress, Daejeon, South Korea, September 25 - 30 2016 , as well as R. Gibbens, J. Boschma and C. Sullivan: “Construction and testing of a new aircraft cycloidal propeller,” 13th Lighter-Than-Air Systems Technology Conference, American Institute of Aeronautics and Astronautics, June 1999 and JH Boschma: “Cycloidal propulsion for UAV VTOL applications”, SBIR Topic Number N98-022, Report prepared for the government technical liaison naval air warfare center - aircraft division, by Bosch Aerospace, November 1998 .

In addition, the flight cabin does not need to be tilted to change the direction of flight and also to fly forward quickly of a cyclorotor-powered aircraft, which offers significant comfort advantages compared to conventional rotary-wing aircraft.

The historical patents US 1656492 A , US 2507657 A , US 2633311 A , US 5265827 A and US 6932296 B2 have specified various versions of aircraft with cyclorotor propulsion.

From the EP 3397551 B1 An aircraft is known which is basically designed in the manner of a conventional rotary-wing aircraft, with a tail rotor being dispensed with and instead two additional cyclorotors being attached to the side for torque compensation and horizontal thrust. The resulting advantage of increased horizontal flight speed is offset by a loss of energy efficiency due to replacing the tail rotor with two cyclorotors, as well as a resulting higher drive requirement as a result of additional resistance and increased structural mass. In addition, the interaction of the rotor wake (the main rotor) can lead to interaction and thus aerodynamic disturbances of the cyclorotors.

From the EP 3715249 A1 An aircraft is known which is basically designed in the manner of a wing aircraft, with four cyclorotors embedded in the wing surfaces and thus ensure propulsion in horizontal flight and provide lift during take-off and landing. The space savings through the integration of the cyclorotors into the wing surfaces are counteracted by the resulting disruption of the flow on the wings. Due to the arrangement of the rotor pairs in a line parallel to the fuselage of the aircraft, an aerodynamic disturbance of the rear rotors due to the vortices from the front rotors is to be expected. In addition, the use of four rotors is expected to result in increased structural weight and thus increased drive requirements compared to conventional rotary wing aircraft.

The DE 102018112079 A1 discloses a vertical take-off rotary wing aircraft which is suitable for carrying loads. By using a conventional helicopter as an aircraft, the disadvantages of helicopters already mentioned, such as the relatively slow response to control inputs and the suboptimal energy efficiency, can be seen as negative in this concept. In addition, the use of a tail rotor (or other drives) to compensate for the yaw moment leads to energy losses, as the energy used cannot be actively used for propulsion or lift.

Another historical patent on the subject of “crane helicopters” with a conventional helicopter is the DE 19950405 A1 to mention, although the same disadvantages of a helicopter occur here.

The historical patent specification US 4482110 A describes an airship-cyclorotor combination for the application of a crane helicopter. The individual rotor blades rotate around the outer skin of the airship. The disadvantages of this application compared to the new invention to be reported are the high complexity of the device and the high expected interference resistance between the rotor blades and the outer skin of the airship. In addition, the airship's large space requirement is a significant disadvantage.

PROBLEM

The concept of conventional helicopters according to the current state of the art suffers from a high level of control complexity and from power losses due to an unavoidable drive (tail rotor, exhaust nozzles, coaxial rotor or similar) for torque compensation, as well as the constant flow interaction between the wake and the cabin or payload.

Additionally, VTOL aircraft are relatively energy inefficient when hovering due to their small rotors or turbines.

The main aim of this invention is to enable an improved figure of merit of the rotor or the entire rotor assembly for the transport of loads or people in slow forward and backward flight, as well as in hovering flight.

This is intended to achieve the goal of being able to transport payloads with the lowest possible power requirement. The primary areas of application of the aircraft are the transport of loads and use as an aviation sports device, as well as the transport of people and use as a cargo crane.

SOLUTION

This problem is solved by the features listed in claim 1. The claim describes, without being limited to, that the aircraft has a single rotor or that all rotors of the aircraft have the same direction of rotation and are either mechanically fixed or coupled by a gear, with the rotor blades on a closed path around the axis of rotation of the Rotors are rotatable and the axes of the wingspan remain parallel to the axis of rotation.

ADVANTAGES ACHIEVED

Based on the aircraft concepts known from the prior art, the invention differs mainly in that no drive is required for the aircraft to generate a counter-torque, both about the transverse and longitudinal axes. The invention thus leads to considerable power savings, since additional drives can be dispensed with, which thus increases the energy efficiency of the aircraft.

A significant advantage is that, in contrast to crane helicopters, the flow wake of this invention does not put pressure on the cabin or the load. The position of the rotor or the rotor assembly also makes it easier to reach the load's destination because the rotor and hull do not have to be above the load. Since the rotor is positioned behind the pilot, a better field of vision for the pilot can be expected compared to multicopters. When the device is controlled by an operator from the ground without a pilot, it is possible for the operator to stand to the side of the device and thus achieve a field of view with little (/no) obstruction.

The typical advantages of cyclorotors can be expected for cyclorotors or similar rotors with rotor blades that rotate parallel to the axis of rotation of the rotor and whose axis of rotation is almost horizontal to the ground. One advantage is the faster response to control commands, as the aircraft can fly forwards and backwards without having to tilt the rotor. In addition, operation is quieter than with conventional rotary-wing aircraft.

In comparison with other cyclorotor aircraft, the required propulsion capacity is reduced by avoiding additional actuators compared to current cyclorotor concepts Technology that allows longer flight times to be achieved with the aircraft and higher payloads can be transported.

A further advantage applies to embodiments with two or more rotors, since the invention is based on the entire aircraft being operated via a single drive unit, preferably by means of a motor (and additional servo motors). This reduces the structural mass of the aircraft relative to conventional cyclorotor aircraft and thus increases energy efficiency.

Since the cyclorotor is located above the center of gravity of the aircraft and the rotor can always point the thrust vector upwards regardless of the pitch angle of the cell, the aircraft is stabilized by the restoring counter torques of the rotor. Stable flight behavior can also be assumed for roll movements, since the rotor is also located above the aircraft and the thrust distribution between the two controllable rotors can be changed without shifting the rotors and thus a roll movement can be controlled.

In the single rotor embodiment, significant weight savings can be expected compared to conventional cyclorotor aircraft while still ensuring stabilization around the transverse and roll axes.

In the embodiments of a single rotor or a firmly connected rotor assembly, the drive power is applied directly to the rotor axis, which offers the advantage of low transmission losses.

Another advantage is the simplicity of the aircraft, which can achieve a variety of different flight states with four control inputs, while the need for counter-torques hardly varies.

BRIEF DESCRIPTION OF THE FIGURES

• 1 zeigt das Fluggerät mit Nutzlastaufnahmehaken;
• 2 zeigt den einzelnen Rotor ohne Steuerungsmechanismus;
• 3 zeigt den Rotorverband mit zwei Steuerungsmechanismen;
• 4 zeigt die Achse des Rotorverbands ohne Rotorblätter;
• 5 zeigt eine Seitenansicht des Steuerungsmechanismus mit einem Rotor;
• 6 zeigt das Strömungsfeld insbesondere den Nachlauf, wobei die Stromlinien die Richtung der Geschwindigkeitsvektoren und die Graustufen die Größe der Geschwindigkeit (dunkler bedeutet schneller) zeigen;
• 7 zeigt in einer Momentaufnahme eine Seitenansicht des Fluggeräts, wobei sich das Fluggerät mit voranschreitender Zeit langsam nach links bewegen würde;
• 8 zeigt den Gelenkarmroboter-Mechanismus (2-RRR Mechanismus);
• 9 zeigt ein Ausführungsbeispiel des Muskelkraft-Fluggeräts;
• 10 zeigt ein Ausführungsbeispiel des Fluggeräts mit drei Rotoren;
• 11 zeigt den Steuermechanismus des mittleren Rotors des Fluggeräts mit drei Rotoren (die Stützprofile, Steuerstange und Blätter des kleinen Rotors werden teilweise verdeckt);
• 12 zeigt das Optimierungsgebiet und die Begrenzung der maximal erreichbaren Position des Gelenkarm-Anbindungspunktes an der Steuerwelle; und
• 13 zeigt eine Seitenansicht des 2-RRR Mechanismus und dessen veränderliche Parameter.

The preferred exemplary embodiments of the present invention are described below with reference to the following figures.

• 1 shows the aircraft with payload hook;
• 2 shows the single rotor without control mechanism;
• 3 shows the rotor assembly with two control mechanisms;
• 4 shows the axis of the rotor assembly without rotor blades;
• 5 shows a side view of the control mechanism with a rotor;
• 6 shows the flow field in particular the wake, with the streamlines showing the direction of the velocity vectors and the grayscale showing the magnitude of the velocity (darker means faster);
• 7 shows a side view of the aircraft in a snapshot, with the aircraft slowly moving to the left as time progresses;
• 8th shows the articulated arm robot mechanism (2-RRR mechanism);
• 9 shows an exemplary embodiment of the muscle-powered aircraft;
• 10 shows an embodiment of the aircraft with three rotors;
• 11 shows the control mechanism of the center rotor of the three-rotor aircraft (the support profiles, control rod and blades of the small rotor are partially obscured);
• 12 shows the optimization area and the limitation of the maximum achievable position of the articulated arm connection point on the control shaft; and
• 13 shows a side view of the 2-RRR mechanism and its variable parameters.

MAIN FORM

The preferred form is in 1 shown and is in the 2 until 8th , 12 and 13 presented in detail. It has two cyclorotors (30) that form an assembly (25); four ser of motors (20), two of which control a rotor; a fuselage (6) which consists of all components of the aircraft that do not rotate with the rotor (30), including any flight controllers, payload hooks (8) and the battery (7). In principle, in this embodiment, no tail units or other measures are required to stabilize the aircraft (24) around the vertical axis (54). Since the landing gear of the aircraft is not relevant to the design of the rotor, it was not taken into account in the drawings. The aircraft (24) can be started either via a landing gear, a starting device on the ground or by hand.

The main form of the control mechanism of a rotor is designed with two servo motors, which together can change the thrust vector in both amount and direction. The fuselage (6) and the rotor assembly (25) are mounted at three points: the two ball bearings on the rotor (28), which are located on the port and starboard side, and the engine on the port side (the drive concept can of course can also be mirrored on a plane perpendicular to the rotor axis). The flight controller and all electronic parts are attached to the fuselage (6). The (payload) load, not shown, is held by the payload receiving hook. The rotor ball bearings (28) are designed as thin-walled ball bearings, so that as much space as necessary is available inside the ball bearing for the control shaft (2).

The aircraft (24) can oscillate (or vibrate) about the rotor axis in the flight state of balanced moments (torque of the motor is compensated by the restoring moment of the fuselage (6)). In its main form, this oscillation can be dampened by a controller or by means of a plate via the aerodynamic resistance.

A core of the invention is the mechanical coupling of several rotors in such a way that several rotors can be operated via a single motor, as in 1 is shown, can be driven. The movement of the motor (21) is transmitted from a gear (22) on the motor shaft via a gear ring (1) to the rotor positioned closest to the motor (21). A drive hub (29) applies the drive load from the ring gear (1) via the connection point (23) to the support profiles (10), which transfer the load to the rotor blade (4) via thin-walled ball bearings at the connection point (11).

Furthermore, the drive load is coupled between the rotors by means of a connection consisting of ball bearings (5), with two separate ball bearings between each pair of blades. In addition, the drive load can be passed on to the next rotor via the rotor axis (3), the support profiles (10) coupled to the fixed connection point (9), and via the ball bearings (5). Through a separate design of the control unit on both sides of the rotor assembly, which in 5 As described, the drive load is supplied by a motor, whereby both rotors can individually determine their thrust vector independently of each other.

The two individual rotors (30) are arranged coaxially to one another so that they form a rotor assembly (25). Within this rotor assembly, the distances (36) between the wing tips (37) of two rotors in the direction of the rotor axis (3) according to claim 5 are to be considered small, with the aim of preventing the formation of edge vortices on the wings as best as possible.

The control mechanism (32) is in 5 , 8th and 13 pictured. Each rotor is controlled via a so-called 2-RRR mechanism, which is based on the principle of an articulated robot with five arms and has two degrees of freedom. The control rods (14) of the rotor blades are mounted axially offset on the control shaft (2). The control command can be transmitted by a servo motor (20) at each of the two attachment points (17). The movement initiated is transmitted to the control shaft (2) at the connection point (19) via the two articulated arms (16) and (18), with the articulated arms being connected to one another via thin ball bearings (15). The movement of the control shaft (2) is now transmitted via another flat ball bearing to the control rod (14), which can then carry out the pitching movement of the rotor blade (4) via a ball bearing at the attachment point (12). The advantage of this mechanism is that the position of the control rods can be precisely controlled. This means that every point in a plane perpendicular to the rotor axis can be controlled with this mechanism, which is often not possible with other control mechanisms. This makes it possible to specify the thrust vector in any amount and direction using just one main control shaft (2).

The position of the servo motor connection points and the length of the articulated arms of the 2-RRR mechanism can be further refined in an optimization process to avoid singularities in operation and to keep the moments around the servo motor shafts low. For this, the length of the Articulated arms and the horizontal and vertical distance (in a plane perpendicular to the control shaft) between the servo motor connection point and the rotor axis in an optimization process. This method of optimizing the 2-RRR mechanism considers an area (51) and evaluates each point within the area according to desired criteria regarding certain properties (see below) of the articulated arm. This process results in a subarea (52) which has the desired properties. Both areas are limited by an outer boundary (53) within which the control shaft can move.

The position of the outer limit (53) is mainly dependent on the inner diameter of the thin-walled ball bearings (28), the diameter of the ball bearing holder on the fuselage, and the diameter of the control shaft. In other words, the diameter of the circle (53) which forms the outer boundary is equal to the inner diameter of the ball bearing minus twice the thickness of the receiving ring on the fuselage minus the control rod diameter.

1. 1) Der Anbindungspunkt (19) an die Steuerwelle (2) muss jeweils von den Gelenkarmen (18) erreichbar sein
2. 2) Die Winkel θ₁ und θ₂ und γ₁ und γ₂ aus 13 sind beschränkt und dürfen einen Grenzwert nicht überschreiten
3. 3) Die Winkel θ₁ und θ₂ müssen zwischen -60° und 60° liegen
4. 4) Die Winkel γ₁ und γ₂ müssen innerhalb von 90°±45° liegen.

The characteristics of the articulated arm are discussed further below:

1. 1) The connection point (19) to the control shaft (2) must be accessible from the articulated arms (18).
2. 2) The angles θ ₁ and θ ₂ and γ ₁ and γ ₂ 13 are limited and may not exceed a limit
3. 3) The angles θ ₁ and θ ₂ must be between -60° and 60°
4. 4) The angles γ ₁ and γ ₂ must be within 90°±45°.

The permitted values for the angles γ ₁ , γ ₂ , θ ₁ and θ ₂ depend on the expected forces transmitted by the control shaft (2) and on the servo motor properties. In addition, the angles depend on the strength of the articulated arms and must be determined individually for each application.

In this case, the angles θ ₁ and θ ₂ are limited by the length of the articulated arm that is rotated by the servo motors: the smaller θ ₁ and θ ₂ are, the further the servo motors have to rotate. The angles γ ₁ and γ ₂ are limited such that the vertical component of the force on the articulated arm is limited, so that the moments that the servo motor has to deliver remain limited.

In order to position the control shaft exactly at its desired location, the following equations are necessary:

und eine Länge Lp, $$L_p=\sqrt{{\left(A_{sx}-P_x\right)}^2+{\left(A_{sy}-P_y\right)}^2}$$

zu berechnen, die zusammen die Lösung des Winkels θ₁ geben: $${\theta }_1=-\frac{1}{2}\pi +{\xi }_1+\text{arccos}\left(\frac{L_1^2-L_2^2+L_p^2}{2L_1{L}_p}\right).$$

The position search is first started at the coordinate of the location (P _x , P _y ), which, together with the coordinates of a connection point between the servo motor and the articulated arm (17), allows an angle ξ ₁ $${\xi }_1=\text{arctan}\left(-A_{sy}+P_{sy}-A_{sx}+P_x\right),$$

and a length Lp, $$L_p=\sqrt{{\left(A_{sx}-P_x\right)}^2+{\left(A_{sy}-P_y\right)}^2}$$

to calculate which together give the solution of the angle θ ₁ : $${\mathrm{<figure-callout\; id="1"\; label="\theta "\; filenames="DE102022114599A1\_0008.png,DE102022114599A1\_0009.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_1=-\frac{1}{2}\pi +{\mathrm{<figure-callout\; id="1"\; label="\xi "\; filenames="DE102022114599A1\_0008.png,DE102022114599A1\_0009.png"\; state="\{\{state\}\}">\xi </figure-callout>}}_1+\text{arccos}\left(\frac{{\mathrm{<figure-callout\; id="1"\; label="L"\; filenames="DE102022114599A1\_0008.png,DE102022114599A1\_0009.png"\; state="\{\{state\}\}">L</figure-callout>}}_1^2-{\mathrm{<figure-callout\; id="2"\; label="L"\; filenames="DE102022114599A1\_0011.png"\; state="\{\{state\}\}">L</figure-callout>}}_2^2+{\mathrm{<figure-callout\; id="2"\; label="L\; p"\; filenames="DE102022114599A1\_0011.png"\; state="\{\{state\}\}">L</figure-callout>}}_p^{}}{2L_1{L}_p}\right).$$

und die Länge L_pB, $$L_{pB}=\sqrt{{\left(B_{sx}-P_x\right)}^2+{\left(B_{sy}-P_y\right)}^2},$$

die zusammen die Lösung des Winkels θ₂ geben: $${\theta }_2=\frac{1}{2}\pi -{\xi }_2-\text{arccos}\left(\frac{L_3^2-L_4^2+L_{pB}^2}{2L_4{L}_{pB}}\right).$$

A similar calculation gives the angle ξ ₂ , $${\xi }_2=\text{arctan}\left(-b_{sy}+P_y,b_{sx}-P_x\right),$$

and the length L _pB , $$L_{pb}=\sqrt{{\left(b_{sx}-P_x\right)}^2+{\left(b_{sy}-P_y\right)}^2},$$

which together give the solution of the angle θ ₂ : $${\mathrm{<figure-callout\; id="2"\; label="\theta "\; filenames="DE102022114599A1\_0011.png"\; state="\{\{state\}\}">\theta </figure-callout>}}_2=\frac{1}{2}\pi -{\xi }_2-\text{arccos}\left(\frac{{\mathrm{<figure-callout\; id="3"\; label="L"\; filenames="DE102022114599A1\_0019.png"\; state="\{\{state\}\}">L</figure-callout>}}_3^2-{\mathrm{<figure-callout\; id="4"\; label="L"\; filenames="DE102022114599A1\_0011.png"\; state="\{\{state\}\}">L</figure-callout>}}_4^2+{\mathrm{<figure-callout\; id="2"\; label="L\; p\; b"\; filenames="DE102022114599A1\_0011.png"\; state="\{\{state\}\}">L</figure-callout>}}_{pb}^{}}{2L_4{L}_{pb}}\right).$$

Note: Instead of the arctan() function, the arctan2() function is preferred in order to avoid divisions by zero and to be able to use the equation in all quadrants.

Asx, Asy

Koordinaten des linken (wie in 13 gezeigt) Servomotor-Anbindungspunkts (17) des 2-RRR Mechanismus;

Bsx, Bsy

Koordinaten des rechten (wie in 13 gezeigt) Servomotor-Anbindungspunkts (17) des 2-RRR Mechanismus;

L1

Länge des linken Gelenkarms am Servomotor des 2-RRR Mechanismus;

L2

Länge des linken Gelenkarms an der Steuerwelle des 2-RRR Mechanismus;

L3

Länge des rechten Gelenkarms an der Steuerwelle des 2-RRR Mechanismus;

L4

Länge des rechten Gelenkarms am Servomotor des 2-RRR Mechanismus;

Px, Py

x- und y-Koordinaten ((x) und (y) wie in 13 definiert) der gewünschten Position der Steuerwelle;

θ1

erforderlicher Winkel am linken Servomotor, um die gewünschte Position der Steuerwelle zu erreichen; und

θ2

Erforderlicher Winkel am rechten Servomotor, um die gewünschte Position der Steuerwelle zu erreichen.

The variables from the above equations are defined as follows:

Asx, Asy

Coordinates of the left (as in 13 shown) servo motor connection point (17) of the 2-RRR mechanism;

Bsx, Bsy

Coordinates of the right (as in 13 shown) servo motor connection point (17) of the 2-RRR mechanism;

L1

Length of the left articulated arm on the servo motor of the 2-RRR mechanism;

L2

Length of the left articulated arm on the control shaft of the 2-RRR mechanism;

L3

Length of the right articulated arm on the control shaft of the 2-RRR mechanism;

L4

Length of the right articulated arm on the servo motor of the 2-RRR mechanism;

Px, Py

x and y coordinates ((x) and (y) as in 13 defined) the desired position of the control shaft;

θ1

required angle on the left servo motor to achieve the desired position of the control shaft; and

θ2

Required angle on the right servo motor to achieve the desired position of the control shaft.

This procedure can be used for both sides of the rotor or rotor assembly to determine the four necessary control angles of the aircraft. The same procedure is also required and used when optimizing a conventional 2-RRR mechanism.

Only the four servo motors are mainly used to control the rotor state, so that the control is carried out by the movement of the two control shafts within the circular area of the ball bearing receiving structure (60) of the fuselage. Since the aircraft is suitable for the hovering flight condition, starboard, port, front and rear “directions” were defined with reference to this flight condition and specified in the nomenclature. These directions do not have to be considered mandatory flight directions. In its main form, forward flight is more stable because the aerodynamic resistance of the fuselage counteracts the torques in this flight condition. It should be noted, however, that the wake in hover or slow flight does not interact with the cabin, fuselage or structure of the aircraft. This is in 6 to see where the wake is thin and roughly opposite to the direction of the thrust. To further control the aircraft, the rotation speed of the rotor can be adjusted to the takeoff mass, which is mainly influenced by the weight of the payload and the energy source.

The flight maneuvers of the aircraft are explained in the following table: Port side Starboard side Px Py Px Py Yaw O +/- O -/+ Climb O - O - Sink O + O + Nod can only be achieved by controlling the speed and the maximum setting angle of the blades (difficult) Push only accessible by rolling (easy) roll +/- O -/+ O Forward flight + - - - Reverse flight - - + -

Notes on the table: Px and Py are on both sides of the aircraft, port and starboard, respectively, as shown in the 13 defined and can be seen from the outside of the aircraft along the transverse axis (55). If the +/characters are used together with the -/+ characters in two columns of the same row, it means that two directions are possible and that either the first character (before the /) or the second character (after the /) is used (other combinations are not permitted). The flight maneuvers in the table are shown as if the thrust vector were acting against a vector between the rotor axis and the control shaft, so the angle between the thrust vector and the vector between the rotor axis and the control shaft is 180° in the direction of rotation. However, the angle is actually slightly larger, as shown in 6 can be seen (here a deviation of 15° is shown). The angle can also depend on the ground effect, the exact rotor geometry (e.g. on the side window) and the fuselage geometry (e.g. on the fairing) and should therefore also be measured on the device to be manufactured in order to adjust the control parameters. This angle can lead to a change in the expected movement of the aircraft, so that, for example, when a control input is made to launch the aircraft vertically, the rotor generates a thrust vector, for example with an upward inclination of 15° (in the direction of rotation (57) of the rotor).

The given movement of the position of the control shaft (P _x , P _y from each rotor side) shows the required change in direction that is necessary to bring the aircraft into the desired movement. Since the vertical axis (54) can oscillate with the fuselage around the rotor axis, this means for the pilot or the design of a flight controller that the angle of the fuselage relative to the horizontal must be taken into account. The coordinate systems for the yaw, climb, (etc.) movements must always be related to the current orientation of the fuselage and the aircraft. The exception to this is the embodiment that attaches the battery or other load to the control mechanism (2-RRR or otherwise), which results in the pilot/flight controller changing the coordinate system with respect to the vertical (54), horizontal (55 ) and longitudinal axis (56) of the battery or load must be set. In addition, it must also be noted that the yaw, roll and pitch movements are related to the center of the rotor axis, which means that the center of the aircraft and the aircraft axis are from 1 and 7 for these movements will move with a combination of rotation and translation. For movements that involve rotation around the vertical or longitudinal axis (i.e. yaw and roll), the rotor reaction is to be expected with an offset of 90° of the rotor due to the gyro effect. This delayed reaction was already taken into account when the tax table was created. The control commands mentioned (within small aerodynamic deviations, as described above) can therefore be used without further adjustments.

For a prototype, the rotor blades (4), as well as the support profiles (10) and control rods (14) were manufactured using a 3D printing process, which were then attached to a CFRP rotor axle. Other conceivable manufacturing processes for 3D printed components include manufacturing from metal or using injection molding for series products. It is also conceivable that these components could be made from wood or composite materials. The wing profiles of the rotor blades were designed as NACA0024 profiles in the CAD design. In the 3D printing process, the filament made of ABS-like material shrinks, ultimately resulting in a profile similar to NACA0026.

The centrifugal forces, which mainly act on the support profiles and the control rods, must also be taken into account. The oscillation of the angle of attack of the blades leads to changing loads that increase quadratically with the speed of the rotor and linearly with the rotor radius. In order to reduce these loads, it is necessary that the rotor blade is designed to be as light as possible and that its center of gravity lies between the pitch (12) and rotation (11) points. The further theory is in the DE102017011890 adequately described.

For this new invention, the ratio between the distance between the connection points of the control (12) and drive (11) rod and the ratio between the diameter of the thin-walled ball bearing of the control mechanism and the diameter of the ball bearing receiving structure (60) is crucial. On the one hand, a large distance between the connection points of the control (12) and drive (11) rods advantageously leads to lower forces in the control rods. However, a larger distance also results in the need for a larger movement of the control shaft to achieve the same angle of attack function. This in turn disadvantageously leads to a larger, thin-walled ball bearing around the rotor axis.

The dimensions of the prototype are given in the following table, as an example only. Changes in, for example, the target payload and the dimensions of the rotor can change the absolute and relative sizes. Naming or description variable absolute size (mm) relatively large (-) Rotor radius R 85.0 1.00 profile chord c 56.7 0.667 Distance between the connection points of the control (12) and drive (11) rod d 16.9 0.199 Control shaft distance (2) e 10.0 0.118 Span of leaves b 85.0 1.00 Distance between the connection point (11) of the drive rod and the leading edge of the blade 19.8 0.233 Distance of the rotor blade center of gravity from the leading edge of the blade 25.5 0.300 Distance between the connection point (12) of the control rod and the leading edge of the blade 36.7 0.432 Control rod length (14) I 85.6 1.01 Articulated arm on the servo motor (16) 26.2 0.308 Articulated arm on the control shaft (18) 44.1 0.519 Horizontal distance between servo motor shaft (17) and control shaft (2) D _x ( 13 ) 45.4 0.534 Vertical distance between servo motor shaft (17) and control shaft (2) D _y ( 13 ) 6.44 0.076 Control shaft diameter (2) 4.00 0.047 Distance A ( 8th ) 48.5 0.571 Distance C ( 8th ) 20.2 0.238

OTHER FORMS

In a preferred embodiment, the aircraft can also be operated as an aircraft with a muscle power drive due to the energy efficiency gain already mentioned due to the savings in propulsion capacity, as in 9 is shown. The aviation equipment can optionally be supported by an electric drive.

If the drive unit of the rotor (muscle power drive is also conceivable) is not connected directly to the rotor axis or to the rotor axis through a gear, the drive can transfer the drive load to the rotor axis via a chain (44) or a toothed belt, as shown in 9 is shown. In this embodiment, a compensating (pitching) moment is also generated by the drive, with this counter-torque being generated for the muscle power drive by introducing a force into the pedal and being transmitted to the rotor via the chain (44). If the drive is via a chain (44), any connection between the fuselage and the rotor axis must be supported. In the case of muscle power propulsion, the person has to hold on to the trunk structure in order to apply the force to the pedal to be able to. If a motor is used instead of muscle power, the torque can be introduced by firmly connecting the motor to the hull.

In a further embodiment, the aircraft consists of a main rotor in the middle and two rotors on each side of the main rotor, with all rotors being arranged on one axis, as shown in 10 is shown. The main rotor could be designed with a larger diameter than the side rotors. All three rotors are driven by a single motor, although the speed of the main rotor could be adjusted via a gearbox or by a motor controller. Because the task of the main rotor is primarily to generate thrust and less to control the aircraft, the side rotors are controlled by the servo motors and the main rotor is only controlled by varying the speed of the motor.

In a further embodiment, the battery is suspended from the control shaft, which results in the control rod trying to always maintain the same position due to the weight of the battery, even when the hull oscillates. An example of this is in 10 shown. This concept offers the advantage that, due to the high battery mass and the independence of the battery position from the fuselage position, an almost constant control command, which leads to an upward thrust vector. For an embodiment with three rotors, the in 10 and 11 is shown, this approach can be used to specify a constant thrust vector direction for the central central rotor in order to save the use of powerful and energy-intensive servo motors, which means that stable flight behavior with better energy efficiency can still be expected.

In an expanded embodiment, a control mechanism is installed on both sides of the rotor for reasons of stability. The rotor, which would otherwise only be controlled from one side in the preferred embodiment, would experience high bending and torsional moments on the blades, which is why the rotor is supported on both sides by means of this expanded embodiment. Using this extension, the battery can also be attached to the control shaft on both sides of the wing, thereby supporting the rotor from both sides. In addition, additional stiffening can be achieved by means of a cross brace in the rotor axis direction between the two control shafts and the use of additional servo motors can be avoided. Stability is particularly important for the central rotor, as it is supposed to absorb the main loads. This concept of stabilizing the rotor by using two control units per rotor can also be used in the other embodiments.

In a further preferred embodiment, the transport of additional trimming and balancing weights can be dispensed with since the center of gravity of the aircraft is positioned sufficiently far in front of the cyclorotor so that the restoring moment imposed by the payload (43) is always greater than the acting one Torque of the drive (58). This can avoid an unstable flight situation. Avoiding additional weights increases the payload capacity of the aircraft.

In a further preferred embodiment, more than two rotors are arranged axes parallel to one another, so that control of the aircraft about the vertical (54) and longitudinal axis (56) by generating opposite thrust vectors of the individual rotors remains possible, as in the main embodiment. This feature has the advantage that additional yaw stability can be achieved through active control of the aircraft.

There can also be a cabin/fuselage between the coaxially arranged rotors.

Further options as optional forms for controlling the are described below and can also be combined with this form.

The axial displacement of the control rods in the preferred embodiment can be implemented by a direct drive (without a gear ratio) if larger ball bearings are used at the connection point between the control rods and the control shaft. However, the diameter of the ball bearing per control rod would correspond approximately to the inner diameter of the ring gear from the preferred embodiment, which is why this embodiment would be inefficient.

In a further embodiment, which is preferably conceivable for the main rotor, all control rods (14) are located in, or almost in, a plane perpendicular to the rotor axis, with one control rod being firmly connected to the outer control ring (46) (50) and the other control rods are mounted on this control ring (46) (49). This embodiment has the advantage that the engine delivers the power directly to the Rotor axis can apply and does not transfer the power to the ring gear via a gear, as in the preferred embodiment. This means that friction losses can be saved here due to the savings in translation. In principle, three embodiments are conceivable with the outer control ring (46). These differ in the degrees of freedom of control available to the rotor. A distinction is made between one and two degrees of freedom.

For the one degree of freedom embodiment, an example of this mechanism is shown in 11 shown. Here the outer ring is mounted on an inner disk (31), which in turn is mounted on the rotor axis, as shown in 11 is shown. This offers the advantage that no or only one servo motor has to be used to control the inner rotor. Like it in 11 is shown, the rotor axis should not be mounted (48) in the center of the inner control disk (31), but at a distance in order to be able to generate a thrust vector. This inner control disk can be passively controlled by being held in one position by the weight of a connection (45) to the fuselage or a high-mass component (battery (7) or payload (43)). This means that the control rods can rotate around the rotor axis on the outer ring of the ball bearing, but the thrust vector direction of the rotor is fixed.

In this embodiment, only a servo motor is used for the control mechanism of a rotor. In this case, the servo motor can only influence the direction of the thrust vector, while the magnitude of the thrust vector can be controlled via the speed of the motor. Since the control mechanism in this case only has one degree of freedom, the aircraft is no longer so easy to control because the thrust control is carried out via the speed of the engine and this therefore leads to longer control times.

For an embodiment of the control with two degrees of freedom, the inner control disk (31) is no longer designed as a disk, but as a ring that is large enough so that the rotor shaft can be guided through it. This enables thrust vector control, both in magnitude and direction. An x-y or 2-RRR mechanism can be used to precisely position the disk in a plane perpendicular to the rotor axis. The inner control ring (31) is mounted on the outer ring (46), so that the inner control ring (31) does not rotate. A control input to the rotor is possible by moving the inner control ring (31) using a mechanism and thus transmitting the translational movement to the outer control ring (46) via the bearing between the rings.

In the above one and two degrees of freedom embodiments, a control mechanism may be attached to the inner control disk (31) or the inner control ring (31) of the rotor. The control signal can be routed via a U-profile from the control unit (e.g. a 2-RRR mechanism), which has an additional servo motor to control the central rotor.

In a further embodiment of the control mechanism, an x-y mechanism (functional principle of a cross table) is used instead of a 2-RRR mechanism, with the 2-RRR mechanism being the preferred embodiment.

In a special embodiment, the aircraft consists of only one rotor with a control mechanism. The rotor positioned on the aircraft according to the invention makes it possible to move the aircraft in the plane perpendicular to the rotor axis, with any change in flight direction from an angle up to 360° in the plane being able to be carried out by the pitching mechanism. Especially when there is a strong crosswind component or when the center of gravity is at a large distance from the rest position in the direction of the rotor axis, stabilization around the vertical axis (54) cannot be easily guaranteed, which is why a fin or rudder is used, which is positioned in the wake of the rotor can be. In another embodiment with a single rotor, the aircraft can be controlled around the vertical axis (54) using a propeller. Furthermore, an embodiment is conceivable in which the aircraft is stabilized from the ground about the vertical axis (54) using guide ropes. It is also conceivable to stabilize the aircraft around the vertical axis (54) by varying the center of gravity, for example by moving a trim weight or the battery.

In a further embodiment with a single rotor, flexible rotor blades are used, which are either completely flexible or only flexible in the area of the mid-span of the wing. This property makes it possible to control the rotor differently from both sides. This allows a similar effect to be achieved as when coupling two rotors, although only one rotor is used here and the leaves are therefore continuous. When different control commands are entered, the rotor blades are slightly twisted in the span direction. This can be useful, for example, for rotating the aircraft about the vertical axis (54). An advantage of this embodiment is the complete absence of the central edge vortices, which leads to increased energy efficiency.

In a further embodiment, the support profiles (10), which represent the connection between the rotor axis and the blades, are replaced by a disk. This disc could, for example, be made of CFRP, wood or a thin PLA plate. Furthermore, the disc can be designed with recesses for weight savings or as a filled disc for better aerodynamic efficiency. A recess in the middle of the disk and a filled disk at the edges would have the advantage that the air can flow through the rotor in the axial direction, but the formation of wing edge vortices would be reduced. It should be noted that with a non-circular design of the rotor cross section perpendicular to the rotor axis, the disk takes on the shape of the non-circular rotor track. Holes are also provided in the disk, which enable the connection between the rotor blade and the control rod (14).

In the preferred embodiment of the aircraft, the rotor blades move in a circular path (42) around the rotor axis, which is characteristic of the concept of a cyclorotor. In a further embodiment, the rotor blades move on a non-circular path, since the basic principle of the aircraft continues to apply as long as a non-negligible moment (58) is generated around the rotor axis when the rotor rotates and the rotor axis is parallel or almost parallel to the transverse axis (55) of the aircraft. For this reason, the rotor blades can also rotate on elliptical paths instead of traditional circular cyclorotor paths. Dynamically changeable rotor orbits are used in the US8540485B2 and the US11198507B2 describe. Also the US9580171B2 ( EP2858894B1 ) and the GB190906378A describe long loop paths for rotors. Other non-circular closed paths cannot be ruled out for this embodiment of the aircraft. When using non-circular rotor orbits, better aerodynamic performance is expected compared to circular rotor orbits.

Any non-circular rotor track as described in claim 14 is a rotor track which eliminates the parasitic aerodynamic drag of a rotor profile by adjusting the effective angle of attack. In order to achieve this, the local flow direction of the rotor blade must always point perpendicular to the desired rotor thrust vector, with the thrust vector pointing upwards when hovering. This would result in a rounded triangular path, with the rotor blade moving on one side of the triangle at a setting angle of 0° relative to the rotor path and parallel to the thrust direction (e.g. upwards in hover). The rotor blades on the other two sides of the triangle move in the opposite direction to the direction of thrust (downwards for hovering flight). For these two sides of the triangle, the angle of the path (between the legs of the triangle) is chosen such that the local flow direction on the blades is perpendicular to the thrust direction. In addition, the setting angle of the blades is selected such that the desired thrust can be generated. The transition between the three straight sections of the path must be optimized using fluid mechanics simulations in order to reduce high stresses in the structure and large jumps in flow velocity. It should be noted that the straight lines of the triangle could also be slightly curved, for example as a spline-based path, in order to avoid jumps in the speed of the rotor blade, which depend on the shape of the path.

In one embodiment, the payload or large, high-mass components are mounted on the aircraft via a hook (8). Fastening via a rope-like connection is also conceivable, but this would lead to reduced flight stability. Furthermore, the mass could be connected to the fuselage via a fixed connection, as in 10 is shown.

In order to increase stability, a plate can be attached in the radial direction of the rotor axis, which creates aerodynamic resistance, which then dampens oscillatory movements of the fuselage or the payload.

In a preferred embodiment of the aircraft, high-mass components (e.g. engines, fuel, batteries and others...) are to be positioned in such a way that the center of gravity (26) is as close as possible to the center of the rotor axis along the transverse axis (55) and as possible is positioned far away from the rotor axis along the longitudinal axis (56) in order to guarantee stability in flight without a payload.

In applications in which the aircraft performs a flight without (payable) load (idle flight), the rotor moments become lower, which is why a lower overall weight of the fuselage is required to counteract the rotor moments. Nevertheless, in these applications it is necessary that the battery, hull and other loads are positioned so that the necessary moments can be achieved. A mechanical or electrical movement of the center of gravity using, for example, an extendable arm is also conceivable in this case.

In a further preferred embodiment, the center of gravity of the aircraft (26) is variable in a user-defined manner. The position of the payload can be changed vertically and/or horizontally using a cable winch mechanism or similar in order to achieve the desired trim.

In a further preferred embodiment, the center of gravity (26) of the aircraft can be changed by a user-defined variation of the position of massive components of the aircraft (e.g. fuel, battery and others...) and can be adjusted vertically and/or horizontally using a cable winch mechanism or similar can be changed to achieve the desired trim.

The position of the center of gravity (26) is very freely defined in claims 3 and 6, as it depends on many parameters. The main goal is that the angle (α) between a line (35) that passes through the rotor axis and the center of gravity and the perpendicular (34) of the rotor axis is between 30° and 60° in all intended flight states. The reason for this is that lower angles increase the likelihood that the fuselage will be in the wake of the rotor. Since the fuselage would potentially enter an uncontrolled rotation if an angle of 90° was exceeded, increasing the angle beyond the described range leads to a reduction in the safety factor. The distance (D) between the rotor axis (3) and the center of gravity (26) depends on both the torque of the rotor (assembly) and the weight of the fuselage. The torque can be described depending on the performance quality of the rotor (assembly), with an increased performance quality (efficiency) leading to a lower torque. Therefore, an efficient aircraft that has a heavy fuselage could even have a distance (D) of 0.5 times the average distance between the axis (3) of the rotor (30) or rotor assembly (25) and the wing attachment point ( 11) can be described in one level. This average distance between the axis (3) of the rotor (30) and the blade connection point (11) is known as the rotor radius for a cyclorotor. An opposite embodiment would be an aircraft that only carries a small payload and a very long fuselage made, for example, of thin CFRP tubes. In such an embodiment, a distance (D) of up to 200 times the rotor radius would be conceivable.

As described in claim 9, in one embodiment the distance (C) between the center of the connection points (23) of the control rods (14) and the closest connection point (19) should not be greater than the distance (A) between the two connection points (19) of the two articulated arms (18). The reason for this is that the moments on the servo motor shafts (17) that do not act around the shaft axis should be eliminated as much as possible, while at the same time the moments around the axis of the servo motor shaft should also be minimized. To ensure that the resulting force of the control rods (14) acts approximately at the center of the connection points (23), the two articulated arms (18) must counteract both the force and the moments. For this reason, the distance (A) serves to avoid the undesirable non-axial moments mentioned above. The larger the distance (A), the lower the force can be on the articulated arm (18) furthest away from the connection points (23) and thus also on the closest articulated arm (18). If the distance (A) is large enough, the servo motors can be made smaller and a weak servo motor can be used to move the furthest away articulated arm (18).

Further embodiments of the invention may become apparent to those skilled in the art upon consideration of the description and application of the invention disclosed herein. For example, ball bearings in the various embodiments can of course also be replaced by other rolling bearings. Therefore, it is intended that the description and examples be considered as exemplary only, with the true scope and subject matter of the invention being defined by the following claims.

REFERENCE SYMBOL LIST

1

Sprocket, only needed on the side of the motor (21).

2

control shaft

3

Rotor axis

4

rotor blade

5

Connection point between the two rotor blades (4) and the support profile (10) connected by a ball bearing for each rotor blade

6

hull

7

Battery case

8th

Payload pick-up hook

9

Fixed connection point between the rotor axis (3) and the support profiles (10)

10

Support profile, also known as rotor arm and can also serve as a drive arm

11

Connection point between the rotor blade (4) and the support profile (10) connected by a ball bearing

12

Connection point between the rotor blade (4) and the control rod (14) connected by a ball bearing

14

Control rod

15

Bearing connection point between the two articulated arms (16) and (18)

16

The articulated arm closest to the servo motor (20).

17

Servo motor shaft or fixed connection point between the servo motor (20) and the articulated arm (16)

18

The articulated arm closer to the control shaft (2) (this number indicates two articulated arms per control shaft)

19

Bearing connection point between articulated arm (18) and control shaft (2)

20

servo motor

21

Drive electric motor

22

Gear for transmitting the torque from the (main) drive motor (21) to the ring gear (1)

23

Connection point between the control shaft (2) and the control rods (14), each individually connected by a ball bearing

24

aircraft

25

Rotor assembly

26

Center of gravity of the torso

28

Rotor ball bearing that serves as a connection between the rotor and the hull (6) on the drive hub (29).

29

drive hub

30

Cyclogyro rotor, shown without control mechanism (32).

31

Inner control disc or ring

32

2-RRR control mechanism

34

Plumb starting from the rotor axis

35

Connecting line from the rotor axis (3) to the center of gravity (26)

36

Distance (36) within a rotor assembly (25) between the wing tips (37) of two rotors in the direction of the rotor axis (3)

37

Wingtips

38

Wingspan

42

Closed path on which the rotor blades move

43

payload

44

Chain

45

Connecting rod between the outer control ring (46) and the battery (7)

46

Outer control ring

47

Thin-walled ball bearing between the outer control ring (46) and the inner control disk (31)

48

Beared connection between the rotor axis (3) and the inner control disk (31)

49

Beared connection between the outer control disk (46) and the control rod (14) (connection on each control rod of the control disk except for a single connection)

50

Fixed connection between the outer control disk (46) and the control rod (14), whereby the outer control disk (46) is rotated by the control rod (14) around the ball bearing (47) (connection only exists once per control unit)

51

Area to be explored by optimizing

52

Sub-area from the area to be optimized (51) that meets the desired properties

53

Maximum size of the circle within the area to be optimized (51) in which the control shaft (2) can move

54

vertical axis

55

Transverse axis

56

Longitudinal axis

57

Rotor rotation direction

58

The torque generated by the rotor, which is not negligible (it outweighs the other torques of the rotor)

60

Mounting the thin-walled ball bearing on the fuselage

A

Distance between the two connection points (19) of the two articulated arms (18), which establish a connection with the control shaft (2).

As

Position (x,y) of the left connection point (17) of the servo motor (20) and the articulated arm (16) of the right 2-RRR mechanism, where the zero point is the center of the rotor axis (3).

Bs

Position (x,y) of the left connection point (17) of the servo motor (20) and the articulated arm (16) of the right 2-RRR mechanism, where the zero point is the center of the rotor axis (3).

C

Distance between the center of all connection points (23) of the control rods (14) from a control unit and the closest connection point (19) of an articulated arm (18) to the control shaft (2)

D

Distance between the rotor axis (3) and the center of gravity (26) in a plane perpendicular to the rotor axis

Dx

Distance in x direction between the control shaft and the servo motor shaft

Dy

Distance in y direction between the control shaft and the servo motor shaft

Fg

Direction of gravity

L

Direction of the buoyancy force or thrust

P

Desired position (x,y) of the control shaft for calculating the angle of rotation of the servo motor shafts, where the zero point is the center of the rotor axis

X

Longitudinal axis (56), with the arrow pointing in the forward direction

x

Local axis for calculating the rotation angles of the servo motor shafts

Y

Transverse axis (55), with the arrow pointing towards starboard

y

Local axis for calculating the rotation angles of the servo motor shafts

Z

Vertical axis (54), with the arrow pointing towards the ground

α

Angle between a connecting line from the rotor axis (3) to the center of gravity (26) and the perpendicular of the rotor axis.

γ1, γ2

Angle between the two articulated arms of a servo motor

θ1, θ2

Angle of rotation of the servo motor shaft around the z-axis

QUOTES INCLUDED IN THE DESCRIPTION

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

Cited patent literature

• US 1753252 A [0008]
• US 2045233 A [0008]
• US 1656492 A [0015]
• US 2507657 A [0015]
• US 2633311 A [0015]
• US 5265827 A [0015]
• US 6932296 B2 [0015]
• EP 3397551 B1 [0016]
• EP 3715249 A1 [0017]
• DE 102018112079 A1 [0018]
• DE 19950405 A1 [0019]
• US 4482110 A [0020]
• DE 102017011890 [0059]
• US 8540485 B2 [0081]
• US 11198507 B2 [0081]
• US 9580171 B2 [0081]
• EP 2858894 B1 [0081]
• GB 190906378 A [0081]

Non-patent literature cited

• S. J. Kim: “Design, development and flight test of vertical take-off and landing UAV, cyclocopter”, Guggenheim Memorial Lecture, 30th ICAS Congress, Daejeon, South Korea, September 25 - 30 2016 [0013]
• R. Gibbens, J. Boschma and C. Sullivan: “Construction and testing of a new aircraft cycloidal propeller”, 13th Lighter-Than-Air Systems Technology Conference, American Institute of Aeronautics and Astronautics, June 1999 [0013]
• J. H. Boschma: “Cycloidal propulsion for UAV VTOL applications”, SBIR Topic Number N98-022, Report prepared for the government technical liaison naval air warfare center - aircraft division, by Bosch Aerospace, November 1998 [0013]

Claims (16)

Aircraft (24), comprising - a rotor (30), consisting of at least one wing (4), which can be rotated on a closed path (42) about the axis of rotation (3) of the rotor (30), while the axis of the wing is in Span direction (wing transverse axis) remains parallel to the axis of rotation (3); - a pivotable mounting of a rotor blade (4) around the blade connection point (11) on the rotor; - a mechanism for varying the angle of attack of a rotor blade (4) relative to and along the closed path (42); and - an aircraft fuselage (6), comprising payloads or the payload connection point (8, 43); characterized in that - the aircraft (24) has a single rotor (30), - or all rotors of the aircraft (24) have the same direction of rotation (57) and are either mechanically fixed or coupled together by a gear to form a rotor assembly (25 ) to build. Aircraft (24) after Claim 1 , characterized in that the aircraft fuselage (6) is mounted in at least one connection point (28,60) on the rotor (30) or rotor assembly (25) by a direct bearing (28) or a motor (21), so that a relative rotation of the rotor (30) or rotor assembly (25) and aircraft fuselage (6) around the axis (3) of the rotor or rotor assembly is possible. Aircraft (24) according to one of the preceding claims, characterized in that the distance between the center of gravity of the aircraft fuselage (6) and the axis (3) of the rotor (30) or rotor assembly (25) is between 0.5 and 200 times the average distance between the axis (3) of the rotor (30) or rotor assembly (25) and the blade connection point (11) in one plane. Aircraft (24) according to one of the preceding claims, characterized in that in a special embodiment with at least two rotors (30), the rotors are arranged axially to one another on a single rotor axis (3), the rotor axis (3) not necessarily having to be continuous , and the rotor assembly is driven by a single drive. Aircraft (24) according to one of the preceding claims, characterized in that the distance (36) within a rotor assembly (25) between the wing tips (37) of two rotors in the direction of the rotor axis (3) is between 0% and 10% of the wingspan (38 ) amounts. Aircraft (24) according to one of the preceding claims, characterized in that the center of gravity (26) of the aircraft fuselage is positioned below the rotor axis (3) during the stable hovering state in such a way that the angle (α) is in the opposite direction of the rotor rotation direction (57) between a connecting line (35) from the rotor axis (3) to the center of gravity (26) and the perpendicular (34) starting from the rotor axis (3) is between 30° and 60°. Aircraft according to one of the preceding claims, characterized in that the mechanism for varying the wing angle of attack relative to the closed path (42) consists of a control rod (14) and a bearing at the connection point (12) of the wing (4), as well as a mounted one There is a connection between the control rod (14) and the central control mechanism (2 or 46). aircraft Claim 7 , characterized in that each central control mechanism (2 or 46), and thus indirectly also the control rods (14), is controlled by a horizontal articulated robot mechanism with 5 arms (2-RRR) ( 5 ) are controlled, the position of the control shaft (2) being moved by two pairs of articulated arms (16 and 18), each with a servo motor (20). aircraft Claim 8 , characterized in that the distance (C) between the center of the connection points (23) of the control rods (14) and the closest connection point (19) is not greater than the distance (A) between the two connection points (19) of the two Articulated arms (18). Aircraft according to one of the preceding claims, characterized in that the control mechanism for varying the angle of attack relative to the closed path, which can include, among other things, the control rod (14) and the servo motors, is designed separately for each rotor. Aircraft according to one of the preceding claims, characterized in that when using several blades (4) per rotor, adjacent blades (4) of a rotor (30) are identical in shape and the angle between adjacent drive rods (13) within each rotor is equal to. Aircraft (24) according to one of the preceding claims, characterized in that the aircraft has only one rotor (30). Aircraft (24) after Claim 1 until 11 , characterized in that the aircraft comprises two rotors (30) on a single rotor assembly (25). Aircraft according to one of the preceding claims, characterized in that the rotor blades (4) move around the rotor axis (3) on a path which has the shape of a rounded triangle. aircraft Claim 1 until 13 , characterized in that the rotor blades (4) move in a circular path around the rotor axis (3). Aircraft according to one of the preceding claims, characterized in that the rotor(s) is/are a cyclogyro rotor(s) (30).

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