DE102017122359A1 - Aircraft in kite configuration - Google Patents

Abstract

The invention relates to an aircraft (1) in a kite configuration, comprising a fuselage (2), an aerofoil (5) and a tailplane (6) spaced from the wing (5) in the region of the stern (3) of the aircraft (1). arranged tailplane (7), which is optionally formed together with a rudder, wherein the wing (5) has at least twice as wide span (SW) as the tailplane (7), and wherein one or more controllable rudder flaps (8, 8 8, 8, 8, 8) are articulated on each side of the fuselage (2) along a respective hinge line (S, S; S, S, S, S) at the trailing edge of the wing (5), and with at least one drive (16a, 11, 16) for driving the aircraft (1). The invention is characterized in that at least one of said rudder flaps (8, 8, 8, 8, 8, 8) is formed on each side of the fuselage (2) as an elevon, which performs both aileron and elevator functions wherein the geometries of the aircraft (1) are designed such that the quotient c / t is greater than -0.25, where t is the mean depth of the wing (5) and cgilt: where c in the x-direction of the aircraft (1) Measured distance from the center of gravity (20) of the aircraft (1) to the respective hinge line center (SP, SP, SP, SP, SP) of the hinge line (S, S, S, S, S, S) of each of the controllable ones Rudder flaps (8, 8, 8, 8, 8, 8) and SL the hinge line length (SL, SL, SL, SL, SL, SL) of the hinge line associated with the respective rudder flap (8, 8, 8, 8, 8, 8) (S, S, S, S, S, S).

Description

The invention relates to an aircraft in kite configuration according to the preamble of the independent claims.

The kite configuration is the configuration commonly used in aviation. Larger passenger planes, smaller propeller aircraft, gliders, etc. are built in this configuration. Typically, these aircraft include a rudder on the rudder for rotating the aircraft about the vertical axis, elevator on the tailplane coupled tailplane for rotating the aircraft about the transverse axis, and ailerons on the wing intended for rotating the aircraft about the longitudinal axis. For certain maneuvers, several of the rudders are operated simultaneously or in quick succession. For example, when cornering without loss of altitude, many aircraft in kite configuration first build a pitch about the longitudinal axis with the aileron. Then, while maintaining the inclination, the elevator is operated so strongly that no loss of height occurs. The rudder can also be operated to enter and exit the turn.

There have been many attempts in the past to realize alternative control concepts for aircraft built in kite configuration. So far, however, no concepts have become known which have proven to be particularly successful in the described 3-axis control.

It is an object of the present invention to provide an aircraft in kite configuration in which the span of the wing is at least twice as large as the span of the tailplane, with a novel control concept that requires less overall construction effort without the flight performance suffers ,

This object is achieved by the features of the independent claims. According to the invention according to claim 1, at least some of the rudder flaps at the trailing edge of the wings are designed as elevons, which take over both the function of an aileron and of an elevator. Here, at least one rudder flap is formed on each side of the hull as such Elevon. In addition, one or more controllable rudder flaps may be provided at the trailing edge of the wing, which only perform the function of an elevator to control the angle of attack. It is also possible that on each side of the fuselage one or more controllable control valves designed as elevons are provided on the hydrofoil, while only one side of the hydrofoil has an controllable control flap designed as an elevator. Generally speaking, it is not necessary that the elevons and also the pure elevator are each arranged in the same number and / or symmetrically to both sides of the fuselage.

According to the invention, the quotient c _medium / t greater than -0.25, where at t the mean depth of the wing ( 5 ) and for c _medium applies: $$c_{mittel}=\frac{{\displaystyle \underset{i}{\Sigma }\left(c_i\cdot S{L}_i\right)}}{{\displaystyle \underset{i}{\Sigma }S{L}_i}}$$

Here is c _i the distance, measured in the x direction of the aircraft, from the center of mass of the aircraft to the respective hinge line center of the hinge line of each of the two controllable rudder flaps on the wing and SL _i the length of the hinge line associated with each rudder flap. The center of gravity lies in the zero point of the coordinate system. As conventionally agreed in the art, starting from this center of gravity, the x-values in the direction of the nose of the aircraft are positive and negative in the direction of the stern. In the above formula, both the controllable rudder valves designed as elevons and the controllable rudder valves designed as pure elevators are taken into account.

In addition, the object is solved by the features of claim 3. In this embodiment of the invention, one or more of the controllable rudder flaps on the wing are designed as Elevons only on one side of the fuselage, while on the other side of the fuselage one or more of the controllable rudder flaps are designed as pure elevator or be controlled as such. It is also possible that on the side of the elevons or one or more pure elevator are provided. at All these cases also apply the above-mentioned formula, which also includes all controllable control valves, regardless of their execution as Elevon or pure elevator.

An embodiment according to claim 3 is thus, for example, that on the left wing half designed as an Elevon controllable rudder flap is present and on the right wing half serving as a pure elevator controllable rudder flap. Even with such a rudder flap constellation, stable and easily controllable flight conditions can be achieved.

In the aircraft according to the invention, it is possible to dispense with an elevator on the horizontal stabilizer. This embodiment without elevator on the horizontal stabilizer is a particularly preferred embodiment according to the invention. Thus, appropriate control elements, possibly aerodynamically unfavorable linkages, drives or actuators, etc. are dispensed with, which not least weight is saved. Further, for example, the failure of a drive for one or more rudder flaps on the wing or any other failure of one or more rudder flaps are easier compensated than the failure of an elevator on the tailplane.

The mean depth t the wing is here, as known from the prior art, defined as the quotient of the wing area divided by the span of the wing. Here, the fuselage portion disposed between the two wing halves is taken into account in the calculation of the wing area such that each wing half is extended at both the nose and tail edges to the longitudinal axis of the aircraft and the thus enclosed area in the area said trunk portion to the in the said mean depth t inflowing wing area is added.

Elevons are known in principle, they represent a combination of elevator and ailerons. From this combination results in the term "Elevon", which is a portmanteau word from the English words "elevator" (elevator) and "aileron" (aileron). Elevons are used with delta-wingers and flying-wingers, both of which have no kite configuration. So far, however, it has not been possible to realize Elevons in aircraft in kite configuration. This is the first time in the context of this invention by means of the features of claim 1 succeeded.

The controllable rudder flaps on the wing, regardless of their training as Elevon or pure elevator, act at a rash in the positive or negative direction at the same time as flaps, which change in their operation, the wing profile in terms of its resistance and its buoyancy. When actuating the controllable rudder flaps downwards (positive angular deflection) not only the profile is optimized for larger angles of attack, but the aircraft is automatically trimmed in an increased angle of attack. Ideally, it is trimmed exactly in the increased angle of attack for which the profile is optimized.

It is a basic idea of the invention that the center of mass of the aircraft (in the direction of flight, independent of the payload) and the hinge line centers, which are essentially the same as the flap pivot points, are relatively close to one another in the x-direction. In this way, the resulting additional buoyancy force generated by positive displacement, i. Deflection down, the controlled rudder flaps is formed, very close to these hinge line centers or flap fulcrums. On the other hand, if the said distance and thus the said quotient are chosen to be larger, a variation of other parameters, for example the tail lever arm, the size of the tailplane, the stability measure, etc., can not compensate for this and airworthiness no longer exists.

Preferred is the said quotient c _medium / t greater than -0.2, also greater than -0.15 or even greater than -0.1 according to certain embodiments. On the other hand, the said quotient c _medium / t preferably less than 0.2, preferably less than 0.1. A particularly preferred embodiment is characterized in that said quotient is less than zero. With decreasing average distance c _medium at constant average depth t of the wing becomes the quotient c _medium / t ever smaller, ie the center of gravity moves, as seen in the x-direction, closer and closer - at least on average - to a theoretical common hinge line of the controllable rudder flaps.

As stated above, the calculation formula applies both to the case that on the wing on each fuselage side exclusively - one or more - designed as elevons controllable control valves are provided, or whether additionally one or more other controllable rudder flaps articulated as pure elevator at the trailing edge of the wing are. Also applies the above formula for the case of claim 3, namely that one or more elevons are provided on only one side of the fuselage and one or more elevators on the other side of the fuselage. Each of these rudder flaps is considered with its hinge line center, hinge line length, and the distance of that hinge line midpoint to the aircraft's center of gravity (viewed in the x-direction of the aircraft) in the formula.

It has proven to be advantageous in certain cases, if in the case of several juxtaposed elevons on each side of the fuselage, the respective inner elevons can be controlled in maneuvers such that they turn out in a positive elevator rudder more than the respective outer. The advantage of this control is that the inner wing section is then forced earlier to buoyancy failure than the outer, and thus the aircraft at low airspeeds independently reduces the angle of attack, without causing it to tip over a wing half. This flight behavior is often perceived by pilots to be particularly pleasant and safe.

Also, constructions have proven to be very suitable for flying, in each of which an externally controllable rudder flap as Elevon and each an inner controllable rudder flap is designed as exclusive elevator. The inner rudder flaps are then always deflected in the same direction, while the outer rudder flaps can optionally also be deflected in the same direction (and then as - additional - serve elevator) or in opposite directions for rotation about the longitudinal axis of the aircraft (aileron function). The reverse construction (inner rudder valves are Elevons, outer rudder valves serve exclusively as elevator) is possible. Other generally comparable constructions, such as two outer and two middle elevons and two inner exclusive elevators on the wing, are also readily possible. Generally speaking, in certain embodiments it is preferred that some of the controllable rudder flaps on the wing are designed as elevons, while the other controllable rudder flaps are designed as pure elevator or are controlled as a pure elevator or serve as a pure elevator.

In the case of the aircraft according to the invention, the arithmetically determined average curvature of the profiles of the wing in straight flight is advantageously between 1% and 4.5%, preferably between 2% and 4%, particularly preferably between 2.5% and 3.5%. Such an average curvature has proven to be a very good compromise between performance and controllability, for example, in the aircraft shown in the following figures.

It has proven to be advantageous - in particular for an aircraft according to the following figures - if the lift coefficient in straight flight above 0.5, for example above 0.6 or 0.9, and / or that the lift coefficient in straight flight below 2, for example below 1.5, preferably below 1.2.

Furthermore, the surface load of the wing is preferably more than 100 g / dm ² . Even heavy loads are therefore transportable by means of the aircraft. In particular, however, higher flying speeds can also be achieved by means of such a high surface loading.

The maximum glide number of the aircraft according to the invention is more than 10, preferably more than 15.

Particularly preferably, a control in the form of an autopilot is provided, wherein the transverse axis of the aircraft is used as at least one control variable. Here, the position of the aircraft is controlled about the transverse axis.

It has been found to be particularly advantageous if the arithmetically determined mean depth of the valve controllable control valves is greater than 25%, preferably greater than 30%. The aircraft according to the invention therefore particularly preferably has a relatively large arithmetically determined average baffle depth, which has proven to be very advantageous for flight characteristics and controllability. Here, the average baffle depth is calculated as arithmetic mean of all local flap depths, each resulting from a local profile section in the x-direction, here locally determined how far the respective local point of the hinge line of the trailing edge side (end bar) of the wing is removed , and this distance is set in relation to the total tread depth at this point.

The wing extension of the wing is preferably between 5 and 25, more preferably between 8 and 18. Wing extension is a dimensionless index of slenderness of a wing. It is defined as the ratio of the square of the span of the wing to the wing surface or alternatively also as the ratio of the span to the mean depth t of the wing (according to the predicted is the mean depth t of the wing defined as the quotient of wing area divided by the wing span). In the aircraft according to the invention, which is preferably operated with a relatively high lift coefficient, the wing extension in the said areas is advantageous. A wing extension in these areas is relatively large and goes hand in hand with a relatively low maneuverability, which is not necessary in certain applications and therefore not disadvantageous.

The tail lever arm is at least one third of the wing span in the aircraft according to the invention and is therefore chosen to be relatively large. The tail lever arm is defined as the x-direction distance from the neutral point of the wing to the neutral point of the tailplane, divided by the mean depth t of the wing. As known from the literature, the two are neutral points N ₁ and N ₂ those points of the profile where the torque remains nearly constant as the angle of attack increases. The relatively large tail lever arm can be a hindrance in the handling of the aircraft, especially in a version with a span of the wing in the range of about 1 to 3 m. It may therefore be envisaged to divide the length of the fuselage of the aircraft so that the transport box does not become too cumbersome.

Particularly preferably, the controllable rudder flaps extend on the wing, so all Elevons and all possibly additionally existing pure elevator together taken over more than 70% of the span of the wing. An extension of the controllable rudder flaps on the wing of over 80%, such as over 90%, and even over 95% up to 100% of the span of the wing is advantageous in certain embodiments of the aircraft according to the invention. In contrast, conventional ailerons are usually made significantly shorter. In the aircraft according to the invention, an embodiment with a large extension of the controllable control valves facilitates the control and improves the efficiency.

A particularly preferred embodiment of the aircraft according to the invention is characterized in that it has at least one rear and one front drive for driving the aircraft, which are arranged spaced apart in the x-direction. In related embodiments, two rear and two front drives, or two rear drives and a front drive, or a rear drive and two front drives for driving the aircraft are provided.

The aircraft according to the invention is advantageously designed as a vertically take-off aircraft, preferably unmanned aerial vehicle (UAV), drone and / or unmanned aerial vehicle (UAS).

The position of the aircraft is preferably substantially the same during take-off (hovering position) and in cruising flight (horizontal flight), wherein the wing is of rigid design and oriented substantially horizontally both in hover and in horizontal flight.

Preferably, some of the drives for driving the aircraft are pivotable by means of a respective pivoting mechanism between a vertical start position for the hover and a horizontal flight position for the horizontal flight. Thus, drives can be used both for the vertical start (in substantially vertically pivoted drives) and in travel or horizontal flight (in horizontal pivoted position).

Said wing is preferably oriented substantially horizontally both in hover and in horizontal flight and allows aerodynamic horizontal flight. The aircraft further preferably has at least one horizontal-direction rear-viewable drive comprising the rear engine and the rear-engine driven rear propeller with respect to the center of gravity of the aircraft, and at least one horizontal-direction controllable front-drive drive comprising a front engine and a from the front engine driven propeller. Here, the motors are preferably electric motors, and at least a part of the drives is pivotable by means of a respective pivot mechanism between a vertical start position for the hover and a horizontal flight position for the horizontal or cruise. Here, the at least one rear or front drive is arranged and designed such that it applies more than 50% of the necessary for hovering the aircraft hovering flight thrust during operation and is designed hinged for horizontal flight. In contrast, the at least one front or rear drive is arranged and designed so that it applies less than 50% of the necessary for hovering the aircraft hovering thrust and more than 50% of the driving force, preferably 100%, in horizontal flight during operation.

Preferably, the center of mass of the aircraft is located off-center with respect to the rear and front propellers, preferably with the aircraft's center of mass being located closer to those propulsion units for the aircraft which provide more than 50% of the hovering thrust.

The diameter of the propellers of those drives which apply more than 50% of the hovering thrust force is preferably greater than the diameter of the propellers, which apply less than 50% of the hovering thrust, preferably at least 5% larger, more preferably at least 10% larger.

• 1 eine Draufsicht auf eine erste Ausführungsform eines erfindungsgemäßen Luftfahrzeugs in Drachenkonfiguration;
• 2 eine Draufsicht auf eine zweite Ausführungsform eines erfindungsgemäßen Luftfahrzeugs in Drachenkonfiguration;
• 3 eine perspektivische Draufsicht auf eine dritte Ausführungsform eines erfindungsgemäßen Luftfahrzeugs in Drachenkonfiguration im Schwebezustand;
• 4 das Luftfahrzeug gemäß der 3 im Schwebezustand, von der Seite gesehen;
• 5 eine perspektivische Draufsicht auf das Luftfahrzeug gemäß der 3 und 4, nun im Horizontalflug;
• 6 das Luftfahrzeug gemäß der 5 im Horizontalflug, von der Seite gesehen;
• 7 eine Draufsicht auf das im Schwebezustand befindliche Luftfahrzeug der 3 - 6 mit eingezeichnetem Massenschwerpunkt, und
• 8 eine vergrößerte Detailansicht von 7.

The invention will be explained in more detail with reference to figures. Show it:

• 1 a plan view of a first embodiment of an aircraft according to the invention in kite configuration;
• 2 a plan view of a second embodiment of an aircraft according to the invention in kite configuration;
• 3 a perspective top view of a third embodiment of an aircraft according to the invention in dragon configuration in limbo;
• 4 the aircraft according to the 3 in limbo, seen from the side;
• 5 a perspective top view of the aircraft according to the 3 and 4 , now in level flight;
• 6 the aircraft according to the 5 in horizontal flight, seen from the side;
• 7 a plan view of the suspended aircraft of the 3 - 6 with marked center of mass, and
• 8th an enlarged detail view of 7 ,

In the 1 is a first, very sketchy embodiment of an aircraft according to the invention 1 presented in dragon configuration. This has a hull 2 with a tail 3 and a nose 4 on. The rear part of the trunk 2 is as a tail boom 6 formed, on which a tailplane 7 is provided. In the present case (and also in the other embodiments illustrated in the figures) comprises the tailplane 6 no elevator, but only a fixed surface, also called hedge fin. A rudder with fixed fin as well as with or without rudder (not shown) may be provided, which in the present case is not essential to the invention. The horizontal stabilizer can also be combined with a vertical stabilizer in the form of a V-shaped tail. In the context of the top view of 1 All these configurations are possible. The hull 2 may be separable into two or more parts in the x direction, which corresponds to the horizontal flight direction, to a transport of the aircraft 1 to facilitate.

On the nose 4 is a drive 16a for the aircraft 1 provided a motor 17a and a propeller 18a which corresponds to the state of the art.

The aircraft 1 has a arranged in its front area, transverse to the hull 2 extending wings 5 on. Below each wing half and below the nose 4 are standing skids 9a . 9b intended. These three runners 9a . 9b form the landing gear of the aircraft 1 ,

According to the invention, at the trailing edge of the two halves of the wing 5 designed as Elevons, controllable rudder valves 8 ₁ and 8 ₂ arranged. These rudder valves 8 ₁ and 8 ₂ take over both the function of the aileron and the elevator. The rudder valves 8 ₁ and 8 ₂ are each along a hinge line S on the wing 5 hinged, with each hinge line S ₁ respectively. S ₂ a hinge line center SP ₁ or SP ₂ , which is in the geometric center of each hinge line S ₁ respectively. S ₂ lies.

In addition, the wing is 5 a medium depth t assigned as the quotient of the wing area divided by the span SW of the wing 5 is obtained. In the 1 is a medium depth t for illustrative purposes only; the plotted mean depth is in fact a quantity calculated as above and therefore can not be pictorially represented.

Hierbei ist c_i der in x-Richtung des Luftfahrzeugs 1 gemessene Abstand vom Massenschwerpunkt 20 des Luftfahrzeugs 1 zum jeweiligen Scharnierlinienmittelpunkt SP_i (in 1: SP₁ und SP₂ ) der jeweiligen Scharnierlinie S_i (in 1: S₁ und S₂ ) jeder Ruderklappe 8_i (in 1: 8₁ und 8₂ ) und SL_i (in 1: SL₁ und SL₂ ) die Länge der zu der jeweiligen Ruderklappe 8_i zugehörigen Scharnierlinie S_i . Vom Massenschwerpunkt im Nullpunkt des Koordinatensystems verläuft die positive x-Richtung in Richtung der Nase des Luftfahrzeugs und die negative x-Richtung in Richtung des Hecks. Der besagte Abstand in x-Richtung ist auch gleichzeitig der jeweilig kleinste Abstand der besagten Scharnierlinienmittelpunkte SP_i von der Schwerachse 21, die durch den Massenschwerpunkt 20 und senkrecht zur Längsachse 22 des Luftfahrzeugs 1 verläuft. Die Abstände c₁ und c₂ gemäß dem in der 1 dargestellten Ausführungsbeispiel sind in dem gewählten Koordinatensystem als negative Werte definiert.Furthermore, the main focus 20 of the aircraft 1 - which does not change relevantly with a corresponding design interpretation even with a payload - drawn. According to the invention applies now that the quotient c _medium / t is greater than -0.25 (not necessarily from the 1 removable, the - as well as the 2 - intends to introduce only the basic structural and geometric concepts). For the calculation of c _medium applies: $$c_{mittel}=\frac{{\displaystyle \underset{i}{\Sigma }\left(c_i\cdot S{L}_i\right)}}{{\displaystyle \underset{i}{\Sigma }S{L}_i}}$$

Here is c _i in the x direction of the aircraft 1 measured distance from the center of mass 20 of the aircraft 1 to the respective hinge line center SP _i (in 1 : SP ₁ and SP ₂ ) of the respective hinge line S _i (in 1 : S ₁ and S ₂ ) of each rudder flap 8 _i (in 1 : 8 ₁ and 8 ₂ ) and SL _i (in 1 : SL ₁ and SL ₂ ) the length of the respective rudder flap 8 _i associated hinge line S _i , From the center of gravity at the origin of the coordinate system, the positive x-direction is in the direction of the nose of the aircraft and the negative x-direction in the direction of the stern. The said distance in the x-direction is also at the same time the respective smallest distance of the said hinge line centers SP _i from the heavy axis 21 passing through the center of mass 20 and perpendicular to the longitudinal axis 22 of the aircraft 1 runs. The distances c ₁ and c ₂ according to the in the 1 illustrated embodiment are defined in the selected coordinate system as negative values.

Gemäß der Erfindung greift auf Grundlage der aerodynamischen Gesetze die resultierende zusätzliche Auftriebskraft, die durch positive, nach unten gerichtete Auslenkung der Ruderklappen entsteht, demnach sehr nahe am jeweiligen Scharnierlinienmittelpunkt bzw. Klappendrehpunkt an. Es hat sich herausgestellt, dass auf diese Weise auch Elevons bei Luftfahrzeugen in Drachenkonfiguration eingesetzt werden können, was in der Vergangenheit nicht gelang.In the case of the first embodiment of 1 after all this, then: $${\text{c}}_{\text{medium}}=\left({\text{c}}_1\times {\text{SL}}_1+{\text{c}}_2\times {\text{SL}}_2\right)\text{/}\left({\text{SL}}_1+{\text{SL}}_2\right),$$

According to the invention, based on the aerodynamic laws, the resulting additional buoyant force resulting from positive, downward deflection of the rudder flaps thus approaches very close to the respective hinge line center or flap pivot point. It has been found that elevons can also be used on aircraft in a kite configuration, something that was not possible in the past.

The said quotient c _medium / t is even preferably greater than -0.2, and may even be greater than -0.15 or greater than -0.1. On the other hand, it has been found that it is advantageous if the said quotient c _medium / t is selected to be less than 0.2, preferably even less than 0.1 or even less than 0. In particular, in wing halves with opposite Flügelpfeilungen in which controllable control valves are provided at each of the differently swept sections, positive quotients c _medium / t be advantageous in terms of flight characteristics. Positive quotients mean large horizontal stabilizers and / or long tail conveyors, which, however, in certain cases can affect the flight performance adversely.

The arithmetically determined mean damper depth of the controllable rudder valves 8 ₁ . 8 ₂ - generally 8 _i - Is particularly preferably greater than 25%, preferably greater than 30%, and thus relatively large compared to conventional flap depths of aircraft in kite configuration. A large average arithmetic mean valve depth is associated with a low quotient c _medium / t ,

Like that 1 and 2 (And the other figures) can be seen, extend the controllable control valves 8 _i in their entirety over much of the span SW of the wing 5 , in each case over more than 90% of the span SW of the wing 5 ,

The embodiment according to the 1 can also be designed such that, for example, the rudder flap 8 ₁ is designed as Elevon and the rudder flap 8 ₂ as a pure elevator (or vice versa). It has been found that even in this constellation, which realizes the features of claim 3, a flight and controllable aircraft can be realized. It also requires in this case no other elevator in the tailplane.

• - Die arithmetisch ermittelte mittlere Wölbung der Profile des Tragflügels 5 im Geradeausflug liegt vorzugsweise zwischen 1% und 4,5%, weiterhin bevorzugt zwischen 2% und 4%, besonders bevorzugt zwischen 2,5% und 3,5%.
• - Der Auftriebsbeiwert im Geradeausflug liegt oberhalb von 0,5, beispielsweise oberhalb von 0,6 oder 0,9.
• - Der Auftriebsbeiwert im Geradeausflug liegt unterhalb von 2, beispielsweise unterhalb von 1,5, vorzugsweise unterhalb von 1,2.
• - Die Flächenbelastung des Tragflügels 5 liegt vorzugsweise oberhalb von 100 g/dm².
• - Die Flügelstreckung beträgt vorzugsweise zwischen 5 und 25, weiterhin bevorzugt zwischen 8 und 18.
• - Der Leitwerkshebelarm beträgt vorzugsweise mindestens ein Drittel der Spannweite des Tragflügels 5. Der Leitwerkshebelarm ist vorliegend definiert als in x-Richtung verlaufender Abstand A (s. 1 und 2) vom Neutralpunkt N₁ der Tragfläche 5 bis zum Neutralpunkt N₂ des Höhenleitwerks 7, geteilt durch die mittlere Tiefe t des Tragflügels 5. Die beiden Neutralpunkte N₁ und N₂ sind hierbei wie üblich als Punkt des Tragflügels 5 bzw. des Höhenleitwerks 7 definiert, in dem das Drehmoment bei einer Anstellwinkeländerung jeweils nahezu konstant bleibt.

The following values are advantageous for the flight characteristics of the aircraft according to the invention 1 according to all here in the 1-8 featured embodiments have proven:

• - The arithmetic determined mean curvature of the profiles of the wing 5 in straight-ahead flight is preferably between 1% and 4.5%, more preferably between 2% and 4%, particularly preferably between 2.5% and 3.5%.
• The lift coefficient in straight-ahead flight is above 0.5, for example above 0.6 or 0.9.
• - The lift coefficient in straight flight is below 2, for example below 1.5, preferably below 1.2.
• - The wing loading of the wing 5 is preferably above 100 g / dm ² .
• The wing extension is preferably between 5 and 25, more preferably between 8 and 18.
• - The tail arm is preferably at least one third of the span of the wing 5 , The tail lever arm is defined herein as extending in the x-direction distance A (S. 1 and 2 ) from the neutral point N ₁ the wing 5 to the neutral point N ₂ of the horizontal stabilizer 7 , divided by the mean depth t of the wing 5 , The two neutral points N ₁ and N ₂ are here as usual as the point of the wing 5 or the tailplane 7 defined, in which the torque remains almost constant at a pitch change.

Erfindungsgemäß sind die Geometrien des Luftfahrzeugs 1 derart ausgelegt, dass der Quotient c_mittel /t größer als -0,25 ist.In the 2 is a second embodiment of an aircraft according to the invention 1 shown in which on each side of the fuselage 2 two elevators designed as elevon control valves in the form of outer rudder valves 8 ₁ and 8 ₂ and two elevators designed as elevons in the form of inner rudder flaps 8 ₃ and 8 ₄ are provided. Each of these rudder valves 8 ₁ . 8 ₂ . 8 ₃ . 8 ₄ has a hinge line S ₁ . S ₂ . S ₃ . S ₄ with a hinge line length SL ₁ . SL ₂ . SL ₃ . SL ₄ and a corresponding geometric hinge line center SP ₁ . SP ₂ . SP ₃ . SP ₄ on. According to the formula of claim 1, an average distance c _medium calculated in the relevant quotient c _medium / t of claim 1 is used. As stated above, this is the average distance c _medium - starting from the center of mass 20 of the aircraft 1 as coordinate zero - calculated from the sum of all products of each hinge line length SL ₁ . SL ₂ . SL ₃ respectively. SL ₄ for the rudder valves 8 ₁ . 8 ₂ . 8 ₃ respectively. 8 ₄ with the associated distance c ₁ . c ₂ . c ₃ respectively. c ₄ (taking only negative values in the present coordinate system), which is measured as the distance of the associated respective hinge line midpoint SP ₁ . SP ₂ . SP ₃ . SP ₄ the relevant rudder flap 8 ₁ . 8 ₂ . 8 ₃ . 8 ₄ to the center of mass 20 , again measured in the x-direction. This sum then becomes the sum of all hinge line lengths SL ₁ . SL ₂ . SL ₃ . SL ₄ divided, the result being the average distance c _medium results. Expressed as a formula: $${\text{c}}_{\text{medium}}=\left({\text{<figure-callout id="1" label="c" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">c</figure-callout>}}_1\times {\text{<figure-callout id="1" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_1+{\text{c}}_2\times {\text{<figure-callout id="2" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_2+{\text{c}}_3\times {\text{<figure-callout id="3" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_3+{\text{c}}_4\times {\text{SL}}_4\right)\text{/}\left({\text{<figure-callout id="1" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_1+{\text{<figure-callout id="2" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_2+{\text{<figure-callout id="3" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_3+{\text{SL}}_4\right)$$

According to the invention, the geometries of the aircraft 1 designed so that the quotient c _medium / t is greater than -0.25.

The calculation according to claim 1 also applies to designs according to the invention with more than two rudder flaps designed as elevons on each wing half. It is also relevant if only a part of the controllable rudder flaps is designed as elevons, while the other part of the controllable rudder flaps on the wing is designed as a pure, i. exclusive, elevator are formed, which are always deflected in the same - positive or negative - direction.

In the 3 - 8th a third embodiment of an aircraft according to the invention is shown, which can start and land vertically. It essentially retains its position in both hovering and horizontal flight. The 3 and 4 show the third embodiment in perspective plan view and in side view in the floating state, while the 5 and 6 the aircraft 1 in travel or horizontal flight represent (horizontal flight direction H , which is the x-direction of the aircraft 1 corresponds).

The aircraft 1 according to the third embodiment has an elongated body 2 with a tail 3 and a nose 4 on. In the illustrated embodiment, the hull is 2 formed in one piece; but it can also consist of several parts, for example, two in the longitudinal direction of the aircraft 1 side by side running trunk parts. The rear part of the trunk 2 is as a tail unit 6 designed with a relatively small diameter at the rear 3 turned end a V-tail 7 (in this case without combined elevator and rudder) is arranged. Also, it is possible that the hull 2 is divided in length, so that it can be accommodated in the form of two shorter body parts easier in a correspondingly short transport box.

In the front of the aircraft 1 is a transverse to the hull 2 extending wing 5 arranged. The wing 5 points away from the longitudinal axis 22 of the aircraft 1 first, a forward pointing negative Flügelpfeilung, while he has then to the wing tips a pointing backwards, positive Flügelpfeilung. At the trailing edge of the wing 5 are inventively controllable rudder valves 8 ₁ . 8 ₂ . 8 ₃ . 8 ₄ intended. Furthermore, below each wing half and below the nose 4 stand skids 9a . 9b intended. These three runners 9a . 9b form the landing gear of the aircraft 1 ,

In the 7 and 8th (The latter is an enlarged detail of the left wing side of the 7 ) are similar to the 2 - The essential for the invention geometric relationships drawn. In principle, the construction resembles the 3-8 in relation to the controllable rudder valves 8 ₁ . 8 ₂ . 8 ₃ . 8 ₄ that of the 2 , Marked in the 7 and 8th especially the two hinge lines S ₁ and S ₂ the outer, in this case designed as Elevons rudder flaps 8 ₁ and 8 ₂ as well as the two hinge lines S ₃ and S ₄ each inner rudder flaps 8 ₃ and 8 ₄ , These can, for example, also act as elevons or as pure elevator. The reverse training is possible, ie the formation of the outer rudder valves 8 ₁ . 8 ₂ as pure elevator and the inner rudder valves 8 ₃ . 8 ₄ as elevons.

The formula according to claim 1 for c _medium applies here in the execution according to the 3-8 also. In the specific case, this is the one above for the second embodiment of the 2 specified formula (which in 8th not shown right wing half is constructed symmetrically to the left, where c ₂ = c ₁ , c ₄ = c ₃ , SL ₂ = SL ₁ , SL ₄ = SL ₃ ): $${\text{c}}_{\text{medium}}=\left({\text{<figure-callout id="1" label="c" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">c</figure-callout>}}_1\times {\text{<figure-callout id="1" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_1+{\text{c}}_2\times {\text{<figure-callout id="2" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_2+{\text{c}}_3\times {\text{<figure-callout id="3" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_3+{\text{c}}_4\times {\text{SL}}_4\right)\text{/}\left({\text{<figure-callout id="1" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_1+{\text{<figure-callout id="2" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_2+{\text{<figure-callout id="3" label="SL" filenames="DE102017122359A1\_0009.png,DE102017122359A1\_0010.png" state="{{state}}">SL</figure-callout>}}_3+{\text{SL}}_4\right)$$

How the particular 8th it can be seen, is the distance c ₃ (and thus the distance c ₄ ) in the chosen coordinate system positive while the distance c ₁ (and thus the distance c ₂ - As well as all distances c _i in the embodiments according to the 1 and 2 ) is negative. Therefore, for the inner controllable rudder valves 8 ₃ . 8 ₄ the sum ( c ₃ × SL ₃ + c ₄ × SL ₄ ) positive. But since the sum ( c ₁ × SL ₁ + c ₂ × SL ₂ ) is negative in terms of outer rudder valves and greater in absolute value c _medium Overall, also negative.

The quotient c _medium / t lies in the embodiment according to the 3-8 at about -0.09. If the sweep of the outer sections of the two wing halves were much more positive, the tailplane significantly larger and / or the tail lever arm much longer, even a positive quotient can result, which can be up to 0.2 according to computational estimates.

On the wing 5 are two pairs of drives 11 . 16 symmetrical to the longitudinal axis of the aircraft 1 arranged. Each of the two wing halves has a rear drive 11 and a front drive 16 on, in each case two of these drives 11 . 16 in the longitudinal direction of the aircraft 1 lie one behind the other. The drives 11 . 16 are each at a rear or front nacelle 10 respectively. 15 arranged, with the gondolas 10 . 15 merge into each other, ie are integrally formed. The rear gondolas 10 are in the direction of the stern 3 Slightly curved upward while the front gondolas 15 essentially horizontally.

At the free end of each rear gondola 10 is a rear drive 11 pivotable by means of a swivel joint 14 arranged, with the rear drives 11 in particular, in each case a rear engine 12 , more preferably an electric motor, and a rear propeller 13 include. The rear propellers 13 are here folded counter to the direction of flight.

Likewise, at the free end of each front nacelle 15 a front drive 16 arranged, with the front drives 16 in particular in each case a front engine 17 , more preferably an electric motor, and a front propeller 18 include.

The to the driving of the drives 11 . 16 other required parts, such as batteries, motor controller or motor plate, etc., are not shown here, however, the skilled person well known. These parts are preferably in the fuselage 2 , the wing 5 or the gondolas 10 . 15 accommodated.

The motors 12 . 17 and propellers 13 . 18 the two drives 11 . 16 are each pivotable by means of a pivot joint 14 . 19 at the gondolas 10 . 15 appropriate. The rear drives 11 are from a downward position ( 3 . 4 ), at which the axis of the engines 12 run vertically, contrary to the direction of flight H swiveling into a horizontal position ( 5 . 6 ), in which the axes of the engines 12 are aligned horizontally. The front drives 16 are from an upward position ( 3 . 4 ), at which the axis of the engines 17 perpendicular, in the direction of flight H swiveling into a horizontal position ( 5 . 6 ), in which the axes of the engines 17 are aligned horizontally.

The two rear propellers 13 have a larger diameter than the front propellers 18 (see also 7 ). Also, with the two rear propellers 13 the quotient of diameter divided by pitch greater than in the front drives, preferably at least 25% larger, more preferably greater by at least 50%, for example by more than 100%. It is also possible and in certain cases preferred that the quotient of diameter divided by slope at the rear propellers 13 150%, for example by 200%, is greater than the front drives 16 ,

Like from the 4 and 6 can be seen, is the center of mass 20 of the aircraft 1 off-center with respect to the rear and front drives 11 . 16 arranged. Although the location of the center of mass is 20 slightly dependent on in particular the pivot states of the drives 11 . 16 , The off-center position of the center of mass 20 but applies to each of these pan states. In the above formula to the average distance c _medium flows the location of the center of mass 20 with drives pivoted for horizontal flight 11 . 16 a (see 5 and 6 ), since the inventive design of the controllable rudder flaps 8 ₁ . 8 ₂ . 8 ₃ . 8 ₄ is aligned to the horizontal flight.

According to the plan view of 7 is the center of mass 20 closer to the rear drives 11 as at the front drives 16 , Here is the distance of the rear drive 11 to the heavy axis 21 passing through the center of mass 20 and perpendicular to the longitudinal axis 22 of the aircraft 1 runs smaller than the distance of the front drive 16 to the heavy axis 21 , The lever arm of the rear drives 11 may preferably be at least 10%, for example more than 20% or more 30%, shorter than the lever arm of the front drives 16 , wherein the lever arms to the shortest distance to the axis of gravity 21 are related. In a specifically implemented embodiment, the distance between the lever arms of the motor axis 12a the front drives 11 to the heavy axis 21 400 mm, while the length of the lever arms from the motor axis 17a the rear drives 16 to the heavy axis 21 at 240 mm. The latter lever arm was thus 40% shorter than the aforementioned lever arm.

In the 3 and 4 is the aircraft 1 shown in the hovering position, in the starting direction S starts and in landing direction L lands. Here are the two rear propellers 13 swung down and push the aircraft 1 in the direction S while the two front propellers 18 are pivoted upwards and the aircraft 1 also in the direction S pull.

It is not mandatory that the aircraft 1 is aligned exactly horizontally in the hover position. For example, it is readily possible for the tail 3 at a small angle, for example, at about 10 °, opposite the nose 4 hangs.

In the 5 and 6 is the aircraft 1 shown in travel or horizontal flight, in which it is in the horizontal direction of flight H flies. Here are the two rear propellers 13 backwards against the direction of flight H panned and tilted while the two front propellers 18 forward in the direction of flight H are pivoted. Only these two front propellers 18 are responsible for the propulsion in horizontal flight in the illustrated embodiment.

The transitions between the different states are as follows: In the 3 and 4 are the rear engines 12 and the front engines 17 in the position in which the aircraft 1 starts. In this state of suspension are the drives 11 down and the drives 16 panned up. For horizontal flight, the drives swing 11 backwards against the direction of flight H and the front drives 16 forward in the direction of flight H , When flying away then fold the two leaves of the two rear propellers 13 due to the wind to the rear against the direction of flight H so the rear propellers 13 do not contribute to propulsion. In the transition from horizontal flight in the limbo state, the front drives 16 pivoted back up while the rear drives 11 restarted and swung down. By this starting the propellers fold 13 due to the centrifugal force again.

The rear drives 11 take over more than 50% of the hovering thrust necessary for hovering and do not contribute to propulsion in horizontal flight. In contrast, take over the front drives 16 less than 50% of the hovering aircraft 1 necessary hover thrust and bring the entire propulsion power in horizontal flight. In the present embodiment, the rear drives 11 therefore not involved in the propulsion. Rather, they fold on or off together due to the wind.

The above task distribution of the rear and front drives 11 . 16 is in the embodiment shown in the figures in particular by the displacement of the center of mass 20 realized. Furthermore, the higher quotient of diameter divided by the slope of the rear drives contributes to an increase in efficiency of the hovering flight. The smaller quotient of diameter divided by the slope of the front drives contributes to an increase in efficiency of horizontal flight.

The control of the drives 11 . 16 , the propeller 13 . 18 , the aileron 8a . 8b etc. can be manual, semi-automatic or automatic. In particular, the transition between the two flight conditions (hovering with front and rear drives switched on - horizontal flying only with front drives switched on) can be realized either manually or automatically. Then, in particular, the propeller speeds are adjusted.

The invention has been described in more detail with reference to an embodiment, but is not limited to this. Variations within the scope of the claims are also possible as a combination of features, even if they are shown and described in different embodiments.

In particular, the aircraft according to the invention can also be controlled by humans and / or used to carry one or more people. In particular, the aircraft according to the invention is suitable as a drone for transporting goods, for controlling and securing the airspace or areas on the ground, for research purposes, for imaging, etc.

LIST OF REFERENCE NUMBERS

1

aircraft

2

hull

3

Rear

4

nose

5

Hydrofoil

6

tail boom

7

tailplane

8 ₁ -8 ₄

controllable rudder valves

9a

Standkufe

9b

Standkufe

10

Gondolas for rear drives

11

rear drives

12

rear engines

13

rear propellers

14

pivot

15

Gondolas for front drives

16

front drives

16a

drive

17

front engines

17a

engine

18

front propellers

18a

propeller

19

pivot

20

Mass center of gravity of the aircraft

21

gravity axis

22

Longitudinal axis of the aircraft

t

mean depth of the wing

c ₁ -c ₄

Distance from hinge line center SP ₁ -SP ₄ to the center of mass 20

c _medium

average distance

S ₁ -S ₄

hinge lines

SP ₁ -SP ₄

Hinge line centers

SL ₁ -SL ₄

Length of hinge lines S ₁ - S ₄

A

distance

N ₁

Neutral point of the wing 5

N ₂

Neutral point of the horizontal stabilizer 7

SW

Span of the wing 5

S

start direction

L

landing direction

H

Horizontal flight direction

Claims (18)

Aircraft (1) in a kite configuration, comprising a fuselage (2), a wing (5), a tailplane (7) spaced from the wing (5) by a tailplane (6) and located in the region of the stern (3) of the aircraft (1) ), which is optionally formed together with a vertical stabilizer, wherein the wing (5) has at least twice as wide span (SW) as the tailplane (7), and wherein one or more controllable rudder flaps (8 ₁ , 8 ₂ ; ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on each side of the fuselage (2) along a respective hinge line (S ₁ , S ₂ , S ₁ , S ₂ , S ₃ , S ₄ ) at the trailing edge of the wing (5) and at least one drive (16a, 11, 16) for driving the aircraft (1), characterized in that at least one of each of the rudder flaps (8 ₁ , 8 ₂ , 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) is formed on each side of the fuselage (2) as an elevon, which takes over both the function of ailerons and elevator, wherein the geometries of the aircraft (1) are designed so that the quotient c _average / t is greater than -0.25, where t is the mean depth of the wing (5) and applies to c _medium : $$c_{mittel}=\frac{{\displaystyle \underset{i}{\Sigma }\left(c_i\cdot S{L}_i\right)}}{{\displaystyle \underset{i}{\Sigma }S{L}_i}}.$$

where c _{i is} the distance measured in the x-direction of the aircraft (1) from the center of gravity (20) of the aircraft (1) to the respective hinge line center (SP ₁ , SP ₂ ; SP ₁ , SP ₂ , SP ₃ , SP ₄ ) of the hinge line (FIG. S ₁ , S ₂ , S ₁ , S ₂ , S ₃ , S ₄ ) each of said controllable rudder flaps (8 ₁ , 8 ₂ , 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) and SL _i the hinge line length (SL ₁ SL ₂ , SL ₁ , SL ₂ , SL ₃ , SL ₄ ) of the hinge line (S ₁ , S ₂ , S ₂ , S ₂ , 8 ₂ , 8 ₂ , 8 ₂ , 8 ₃ , 8 ₄ ) belonging to the respective rudder flap (8 ₁ , 8 ₂ ; ₁ , S ₂ , S ₃ , S ₄ ). Aircraft (1) to Claim 1 , characterized in that only some of said controllable rudder flaps (8 ₁ , 8 ₂ ; 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on the wing (5) are designed as elevons, while the one or more controllable rudder flaps as a pure elevator are formed or controlled as a pure elevator. Aircraft (1) in a kite configuration, comprising a fuselage (2), a wing (5), a tailplane (7) spaced from the wing (5) by a tailplane (6) and located in the region of the stern (3) of the aircraft (1) ), which is optionally formed together with a vertical stabilizer, wherein the wing (5) has at least twice as wide span (SW) as the tailplane (7), and wherein one or more controllable rudder flaps (8 ₁ , 8 ₂ ; ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on each side of the fuselage (2) along a respective hinge line (S ₁ , S ₂ , S ₁ , S ₂ , S ₃ , S ₄ ) at the trailing edge of the wing (5) and at least one drive (16a, 11, 16) for driving the aircraft (1), characterized in that only on one side of the fuselage (2) at least one of said rudder flaps (8 ₁ , 8 ₂ , 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) is designed as an elevon, which assumes both the function of aileron and elevator, and d at least one more of said rudder flaps (8 ₁ , 8 ₂ ; 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on the other side of the fuselage (2) is designed as a pure elevator or is driven as a pure elevator, wherein the geometries of the aircraft (1) are designed such that the quotient c _medium / t is greater than -0.25, where t is the mean depth of the wing (5) and applies to c _medium : $$c_{mittel}=\frac{{\displaystyle \underset{i}{\Sigma }\left(c_i\cdot S{L}_i\right)}}{{\displaystyle \underset{i}{\Sigma }S{L}_i}}.$$

where c _{i is} the distance of the center of gravity (20) of the aircraft (1) measured in the x direction of the aircraft (1) to the respective hinge line center (SP ₁ , SP ₂ ; SP ₁ , SP ₂ , SP ₃ , SP ₄ ) of the hinge line (FIG. S ₁ , S ₂ , S ₁ , S ₂ , S ₃ , S ₄ ) each of said controllable rudder flaps (8 ₁ , 8 ₂ , 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) and SL _i the hinge line length (SL ₁ SL ₂ , SL ₁ , SL ₂ , SL ₃ , SL ₄ ) of the hinge line (S ₁ , S ₂ , S ₂ , S ₂ , 8 ₂ , 8 ₂ , 8 ₂ , 8 ₃ , 8 ₄ ) belonging to the respective rudder flap (8 ₁ , 8 ₂ ; ₁ , S ₂ , S ₃ , S ₄ ). Aircraft (1) according to at least one of the preceding claims, characterized in that said quotient c _average / t is greater than -0.2, preferably greater than -0.15, for example greater than -0.1. Aircraft (1) according to at least one of the preceding claims, characterized in that said quotient c _average / t is less than 0.2, preferably less than 0.1, for example less than 0. Aircraft (1) according to at least one of the preceding claims, characterized in that no elevator is provided on the horizontal stabilizer (7). Aircraft (1) according to at least one of the preceding claims, characterized in that the arithmetically determined mean curvature of the profiles of the wing (5) in straight flight between 1% and 4.5%, preferably between 2% and 4%, more preferably between 2.5% and 3.5%. Aircraft (1) according to at least one of the preceding claims, characterized in that the lift coefficient of the aircraft (1) in straight flight is above 0.5, for example above 0.6 or 0.9, and / or that the lift coefficient in straight flight is below 2, for example below 1.5, preferably below 1.2. Aircraft (1) according to at least one of the preceding claims, characterized in that the surface loading of the wing (5) is above 100 g / dm ² . Aircraft (1) according to at least one of the preceding claims, characterized in that the arithmetically determined average flap depth of said controllable rudder flaps (8 ₁ , 8 ₂ ; 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on the wing is greater than 25%, preferably greater than 30%. Aircraft (1) according to at least one of the preceding claims, characterized in that the wing extension of the support surface (5) is between 5 and 25, preferably between 8 and 18. Aircraft (1) according to at least one of the preceding claims, characterized in that the tail arm lever is at least one third of the span (SW) of the wing (5). Aircraft (1) according to at least one of the preceding claims, characterized in that the controllable rudder flaps (8 ₁ , 8 ₂ , 8 ₁ , 8 ₂ , 8 ₃ , 8 ₄ ) on the wing (5) over more than 70%, preferably over 80%, such as over 90%, of the span (SW) of the wing (5). Aircraft (1) according to at least one of the preceding claims, characterized in that it comprises at least one rear and one front drive (11, 16) for driving the aircraft (1), which are arranged spaced apart in the x-direction. Aircraft (1) according to at least one of the preceding claims, characterized in that it is designed as a vertically take-off aircraft (1), preferably unmanned aerial vehicle (UAV), drone and / or unmanned aerial vehicle (UAS). Aircraft (1) according to at least one of the preceding claims, characterized in that the position of the aircraft (1) during take-off, ie in the hover position, and in cruising, ie in level flight, is substantially the same, the wing (5) rigidly formed and oriented substantially horizontally both in hovering and in horizontal flight, preferably at least some of the drives (11, 16) for driving the aircraft (1) by means of a respective pivoting mechanism between a vertical start position for the hover and a horizontal flight position for horizontal flight are pivotable. Aircraft (1) according to at least one of the preceding claims, characterized in that two rear and two front drives (11, 16) for driving the aircraft (1), or two rear drives and a front drive for driving the aircraft, or a rear Drive and two front drives for driving the aircraft (1) are provided. Aircraft (1) according to at least one of the preceding claims, characterized in that said wing (5) is oriented substantially horizontally in both hover and horizontal flight and enables aerodynamic horizontal flight, the aircraft (1) continuing to move in relation to the aircraft Center of gravity (20) of the aircraft (1) at least one seen in horizontal flight direction rear adjustable drive (11), comprising a rear motor (12) and a rear motor (12) driven rear propeller (13), and at least one seen in the horizontal direction of flight front variable speed drive (16) comprising a front motor (17) and a propeller (18) driven by the front motor (17), the motors (12, 17) preferably being electric motors, and at least a portion of the drives (11, 16 ) by means of a respective pivot mechanism between a vertical start position for the hover and a horizontal flight posititi can be pivoted for the horizontal or cruise, and wherein the at least one rear or front drive (11, 16) is arranged and designed so that it applies more than 50% of hovering force required to levitate the aircraft (1) during operation and is designed hinged for horizontal flight, while the at least one front or rear drive (16, 11) is arranged and designed such that it is less than 50% of the hovering force required to levitate the aircraft (1) and more than 50% during operation % of the driving force, preferably 100%, applies in horizontal flight.

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