Vertical takeoff and landing aircraft

Abstract

The present invention provides a vertical take-off and landing aircraft, comprising a fuselage (2) for transporting passengers and / or loads; a front wing (3) attached to the fuselage; a rear wing (4) attached to the fuselage, the rear wing being behind the front wing in the forward flight direction; a right connecting beam (5a) and a left connecting beam (5b), the connecting beams (5a, 5b) being structurally connected to the front wing and the rear wing, the connecting beams being spaced apart from the fuselage; and at least two lifting units (M1-M6) on each connecting beam, each lifting unit comprising at least one propeller (6b) and at least one motor (6a) driving the propeller, preferably an electric motor, and the lifting units being arranged such that their respective propeller axes are in a substantially vertical direction (z); at least part of the front wing has a sweep angle γ, 45° ≤ γ ≤ 135°, preferably γ = 75°, and at least part of the rear wing has a forward sweep with a sweep angle β ≥ 30°, preferably β = 65°.

Description

Technical Field

[0001] The present invention relates to a vertical take-off and landing (VTOL) aircraft, in particular an aircraft with electric propulsion or lift units, i.e., a so-called electric vertical take-off and landing aircraft. Background Art

[0002] Some similar aircraft designs have been disclosed, such as in USD311720S, US3834654A, and US9499266B1.

[0003] Generally, contemporary electric vertical take-off and landing aircraft designs have poor aerodynamic efficiency because they combine multiple systems to vertically lift the aircraft and move it horizontally. In particular, no known design can be considered a naturally stable aircraft. Examples of known designs include the aircraft of Zee Aero / Cora, see US2018 / 105268A1; the aircraft of EmbraerX; the aircraft of Aurora Flight Science (now part of Boeing); the aircraft of CityAirbus&Project Zero; the aircraft of AutoflightX, see US10081436B1.

[0004] The Pipistrel Nuuva V300 is an unmanned cargo aircraft with its main weight located centrally, which means that the wing surface distribution is basically symmetric about the y-z plane. It has a tandem wing and eight lift propellers. The disadvantage is that the landing load (similar to the Boeing / Aurora PAV) is transmitted through the engine nacelles, which is unfavorable.

[0005] No known design enables passengers to enter the fuselage comfortably. All known designs result in solutions that are unfavorable to passengers entering the aircraft. This is mainly due to the low vertical position of the propulsion and lift systems.

[0006] Most known designs have high aerodynamic blockage (or installation losses) of aircraft components within the lift propeller airflow during hover / low-speed conditions (see the UBER elevate crm003 model), which results in high power losses during hover and transition flights.

[0007] Most known designs have lift propellers behind each other. The more there are, the higher the aerodynamic losses of the rear propellers, which leads to (unnecessary) high power consumption, thereby increasing the overall weight of the aircraft or shortening the range.

[0008] Most known designs are not scalable in size and have a span greater than 10 m, with a relatively low payload capacity. Therefore, scaling up to higher payloads and / or passenger capacities would result in larger aircraft sizes. However, the size of the landing sites is limited, especially in urban areas, and with increasing size, the weight increases disproportionately.

[0009] Most known configurations have many small-diameter lift propellers and thus have a high disk load, which leads to an increased noise level.

[0010] Known designs lack an aircraft architecture that is structurally integrated and results in high weight penalties or aircraft with a low payload-to-empty weight ratio.

[0011] Known designs lack a structurally integrated aircraft architecture, which results in high weight penalties or aircraft with a low payload-to-empty weight ratio.

[0012] In a previous application (EP 3 757 004 A1), the applicant proposed a vertical take-off and landing aircraft design that overcomes most, if not all, of the above disadvantages.

[0013] However, to provide a reliable daily vertical take-off and landing aircraft, an improved hybrid or transitional aircraft (i.e., multi-rotor + winged aircraft) is additionally required in order to:

[0014] - Improve aerodynamic efficiency and / or minimize drag, especially when using a multi-rotor tandem wing, and be aerodynamically efficient in terms of lift-to-drag ratio;

[0015] - Be aerodynamically naturally stable as an aircraft and thus generate an overall aircraft momentum to stabilize the aircraft about its longitudinal axis. This is crucial for more easily meeting certification requirements and improving the overall flight efficiency of the aircraft;

[0016] - Create an aircraft wing configuration that is equally applicable in multi-rotor and flight modes, taking into account that in each case the forces and momenta of the rotating lift units (propellers) and the static lift units (wings) vary greatly and can lead to adverse aerodynamic characteristics;

[0017] - Cope with at least one engine or propeller failure, which means considering the specific center-of-gravity position in relation to the center of drive. Summary of the Invention

[0018] The object of the present invention is to propose a novel design of a vertical take-off and landing aircraft, which provides a further improvement over the design described in EP 3 757004 A1, thereby realizing a vertical take-off and landing aircraft that has increased natural stability, increased aerodynamic efficiency, and equally good aerodynamic characteristics in multi-rotor and flight modes and can cope with the situation of at least one engine or propeller failure (one engine inoperative - OEI).

[0019] This object is achieved by a vertical take-off and landing aircraft and defines a favorable further development of such an aircraft.

[0020] According to the present invention, a vertical take-off and landing aircraft comprises: a fuselage for transporting passengers and / or a load; a front wing attached to the fuselage; a rear wing attached to the fuselage, the rear wing being behind the front wing in the forward flight direction; a right connecting beam and a left connecting beam, the connecting beams being structurally connected to the front wing and the rear wing, the connecting beams being spaced apart from the fuselage; and at least two lifting units on each of the connecting beams, each lifting unit comprising at least one propeller and at least one motor driving the propeller, preferably an electric motor, and the lifting units being arranged such that their respective propeller axes are in a substantially vertical direction; wherein at least part of the front wing has a sweep angle γ, 45° ≤ γ ≤ 135°, preferably γ = 75°, and at least part of the rear wing has a forward sweep (i.e., β < 90°) with a sweep angle β ≥ 30°, preferably β = 65°.

[0021] A "swept wing" means a wing that is inclined backward or forward from its root rather than straight to the side (90°) direction. The term "swept wing" is usually used to denote "rear sweep", but variants include forward sweep, variable sweep, and a skew wing with forward sweep on one side and rear sweep on the other side. In this document, a sweep angle less than 90° indicates forward sweep, i.e., the wing is inclined forward in the (forward) flight direction.

[0022] In the context of the present invention, the sweep angle γ of the front wing can be 45° ≤ γ ≤ 135°. Preferably γ = 75°. The (forward) sweep angle β of the rear wing can be 30° ≤ β ≤ 90°. Preferably β = 65°.

[0023] For better accessibility, in a side view of the aircraft, the connecting beams are preferably placed above or over the fuselage.

[0024] More generally, the present invention can include placing the following physically definable points (which are well-known to those skilled in the art of aviation) in specific positions in order to create a conversion aircraft that has good aerodynamic characteristics while having the characteristics of a high-performance multi-rotor aircraft, especially in the case of one engine inoperative (OEI).

[0025] The definitions of the points mentioned above are as follows and the term will be used hereinafter:

[0026] · The Center of Actuation (CoA) is a reference point defined by the position and attitude of the lifting rotors. The CoA is usually equal to the geometric center of all lifting rotors, but mathematically it is the point where all lift and moments are in balance.

[0027] · The Center of Gravity (CoG) is the point at which all gravitational forces act if the aircraft is in equilibrium. In other words: In physics, the center of gravity or centroid of a mass distribution in space (sometimes called the balance point) is the unique point at which the weighted relative position sum of the distributed mass is zero. This point is defined by the aircraft layout, geometry, and weight distribution.

[0028] · The Neutral Point (NP) is the point at which, mathematically, all moments acting on the aircraft by any lifting surface (such as a wing) are zero. A mathematical analysis of the longitudinal static stability of a complete aircraft (including the horizontal stabilizer) yields the center of gravity position at which the stability is neutral. This position is called the neutral point. (The larger the area of any horizontal stabilizer, the larger the moment arm of the horizontal stabilizer about the aerodynamic center, and the further aft the neutral point.) A well-defined NP is crucial for a longitudinally stable aircraft and is mainly defined by static lifting units such as wings and stabilizers.

[0029] · The Aerodynamic Center (CA) is the point at which the pitching moment (Cm) coefficient of a lifting body (such as a wing) does not vary with the angle of attack (dCm / dα = 0). This point is mainly defined by static lifting units such as wings and stabilizers in terms of plane, position, dimensions, etc. More specifically, the torque or moment acting on a wing through which a fluid flows can be explained by the net lift and net drag applied at a point on the wing and the separate net pitching moment about that point, the magnitude of which varies with the position at which the lift is applied. The CA is the point at which the pitching moment coefficient of the wing does not vary with the lift coefficient (i.e., the angle of attack), making the analysis simpler.

[0030] The present invention proposes to find a technical solution to bring the center of gravity, the center of actuation, and the neutral point as close to each other as possible during the flight phase of transitioning from hover to flight mode and vice versa. In addition to a low-drag design and an integrated airframe design, the layout of the aircraft is preferably designed in such a way that the longitudinal stability defined by Cm / α is negative (i.e., Cm / α < 0).

[0031] When any aircraft moves, it will be subject to minute changes in the forces acting on it and its speed. If such a change causes a further change that tends to restore the aircraft to its original speed and direction without the input of a person or machine, the aircraft can be said to be statically stable. Such an aircraft has positive stability. On the other hand, if such a change causes a further change that tends to make the aircraft deviate from its original speed and direction, the aircraft can be said to be statically unstable. Such an aircraft has negative stability. As is known to those skilled in the art, an aircraft can have a form of low negative stability called relaxed stability to provide ultra-high maneuverability.

[0032] The present invention may include selecting the coordinate positions of the lifting units in such a way that the positions are arranged substantially symmetrically in terms of the geometry of the aircraft, thereby defining the CoA to be geometrically at least approximately at the same x-position as the CoG, while the NP is designed to be located behind these two points (CoA, CoG).

[0033] Generally, the main axes of an aircraft are defined as:

[0034] · Longitudinal axis, or roll axis - the axis that passes through the fuselage of the aircraft from the tail to the nose in the normal flight direction, or the direction the pilot is facing, similar to the waterline of a ship.

[0035] · Transverse axis, or lateral axis, or pitch axis - in a manned aircraft, it extends from the left side of the pilot to the right side and is parallel to the wings of a winged aircraft and parallel to the longitudinal section line.

[0036] · Normal axis, or vertical axis - the vertical axis drawn from top to bottom and perpendicular to the other two axes.

[0037] Hereinafter, these axes are also represented by the letters x, y, and z respectively.

[0038] The present invention may also include a specific balance of the wing surface size (or area) of the aircraft between the front wing and the rear wing, as this defines the longitudinal stability of the aircraft. The above-mentioned forward sweep of these wings pushes the neutral point (NP) of the aircraft as far as possible towards the tail.

[0039] In a corresponding embodiment of the aircraft according to the present invention, the drive center of the aircraft is geometrically at the same longitudinal position as the center of gravity of the aircraft, while the neutral point is located behind the drive center and the center of gravity.

[0040] In summary, the front wing and the front wing have the necessary size (area) to generate the lift required to carry the entire weight of the aircraft.

[0041] However, the planform dimensions of the front wing and the rear wing may not be equal, with the front wing being 40% and the rear wing being 60%, but the lift generated can have the opposite ratio, with the front wing being 60% and the rear wing being 40%. To achieve this, the rear wing can be divided into two separate but structurally connected parts, namely the aircraft lifting part (wing) and the stabilizing unit (e.g., in the form of a so-called V-tail), which contributes little to the lift of the aircraft but provides a significant horizontal surface as a weather vane contributing to lateral stability.

[0042] In a corresponding embodiment of the aircraft according to the invention, the wing area S of the front wing 前 is less than the wing area S of the rear wing 后 , preferably 60% < S 前 / S 后 < 100%, more preferably 60% < S 前 / S 后 < 70%.

[0043] In another embodiment of the aircraft according to the invention, in a plane perpendicular to the longitudinal axis (x) of the aircraft, the rear wing extends laterally from the fuselage such that the angle (θ) formed by the respective half-wings is less than 180°, preferably approximately 90° to 100°.

[0044] To reduce the downwash effect of any rotating lifting unit (e.g., a propeller), which occurs if the airflow generated by the lifting unit impinges on a solid material (e.g., an engine mount), a wing attachment can be placed between the rotor planes. The rotor planes can be defined as the respective areas swept by the rotating propellers. This results in an improvement in rotor performance during hover or low-speed maneuvers.

[0045] Due to the presence of the front wing, the rear wing is generally aerodynamically negatively affected because the front wing causes the airflow to impinge on the rear wing. To reduce this effect, in a corresponding embodiment of the aircraft according to the invention, the rear wing can be placed higher in the vertical position. However, it can still be connected to the connecting beam. Generally, this feature enables a rigid airframe design from the body (fuselage) to the wings and from the wings to the lifting units.

[0046] Since the front wing and the rear wing are structurally connected by a connecting beam, any torsion occurring in the wings can be significantly reduced. This is beneficial because the loads are balanced and routed.

[0047] As described above, the lifting units of the aircraft can be placed substantially symmetrically with respect to the longitudinal (x) and lateral (y) axes of the aircraft. In this way, the coupling in the multi-rotor mode is minimized, and the performance during the OEI mode is improved because the respective geometric distances from each lifting unit to the center of gravity are substantially equal.

[0048] In a corresponding embodiment of the aircraft according to the invention, the lifting units are thus placed symmetrically with respect to the longitudinal axis and the lateral axis of the aircraft, respectively.

[0049] As described above, the position of the lifting units defines the center of action (CoA) of the multi-rotor mode, which can be defined according to the longitudinal position by where x i represents the longitudinal position of the lifting units, and, where n = 6 in the case of a hexacopter configuration.

[0050] In a corresponding embodiment of the aircraft according to the invention, the aircraft includes three lifting units on each connecting beam.

[0051] This also applies to a quadcopter (n = 4) or an octocopter (n = 8) or any other coaxial configuration (with respect to the longitudinal axis of the aircraft).

[0052] To reduce the downwash effect from any (rotating) lifting unit, the static lifting units (i.e., the wings) can be effectively connected to the structure or the connecting beam at a (longitudinal) position (i.e., the area swept by the respective propellers) between the rotating disks or the rotor planes of each rotating lifting unit.

[0053] In a corresponding embodiment of the aircraft according to the invention, the attachments of the front wing and the rear wing, which are respectively connected to the connecting beam, are placed between the respective rotor planes swept by the propellers so that there is no vertical overlap between the propellers and the attachments.

[0054] The aircraft according to the invention has a front wing and a rear wing, which together define the neutral point (NP) of the aircraft. As is known to those skilled in the art, the NP is mainly defined by the combination of the respective aerodynamic centers (CA) of the aircraft wings. The distance between each aerodynamic center (CA 前 for the front wing and CA 后 for the rear wing) multiplied by the longitudinal distance x from the respective center of gravity (CoG) of the aircraft defines the longitudinal momentum. In the context of the present invention, due to the presence of the front wing and the rear wing, the two aerodynamic centers must be considered and balanced. As described above, Cm / α can be selected to be negative, thus obtaining a stable aircraft.

[0055] More specifically, longitudinal static stability is achieved by distributing the horizontal surface area of the aircraft in such a way that a pitch deviation causes a pitching moment to be counteracted. All horizontal surface area located behind the center of gravity contributes positively to longitudinal static stability through the weathervane effect. All horizontal surface area located in front of the CoG is destabilizing. Each horizontal surface area contributes to this effect through its respective area multiplied by the lever arm from its area center to the CoG. The neutral point (NP) of the aircraft is defined as the point at which all pitching moments are in balance. To achieve longitudinal static stability, the neutral point needs to be located behind the CoG. This geometry results in Cm / α < 0.

[0056] Thus, in another embodiment of the aircraft, the longitudinal aircraft stability defined by the quantity Cm / α is negative, where Cm is the pitching moment coefficient and α is the angle of attack.

[0057] Traditional prior art aircraft wings can be attached to the fuselage at a substantially transverse angle (90°), resulting in straight wings (unswept). Following this principle would result in a relatively forward neutral point, thus causing unstable longitudinal stability. To move the aerodynamic center of each wing backward (towards the tail), the forward wing includes a sweep angle γ less than 90°, and the rear wing includes a sweep angle β also less than 90°. As described above, the wing area S of the forward wing 前 is less than the wing area S of the rear wing 后 , but the rear wing surface may generate a smaller vertical lift due to the presence of the V-tail.

[0058] To further reduce the length of the fuselage while maintaining a sufficient rear surface wing area S 后 , the V-tail can include a sweep angle δ greater than 90°. This also improves the load transfer to the fuselage body.

[0059] In a corresponding embodiment of the aircraft according to the present invention, the rear wing portion is configured in the form of a V-tail having a sweep angle (δ) greater than 90°.

[0060] Furthermore, the forward wing can include a dihedral angle (λ) which, in combination with the above-mentioned sweep angle (γ), has a positive effect on the lateral / directional stability of the aircraft.

[0061] In a corresponding embodiment of the aircraft according to the present invention, the forward wing includes a dihedral angle (λ), preferably increasing at least at the respective tips of the forward wing or towards the respective tips of the forward wing, preferably 0° < λ < 5°.

[0062] The rear wing can include an anhedral angle (ε) in order to generate a high vertical offset in the region of the higher downwash of the forward wing, which mainly occurs in the inner part (spanwise direction) of the forward wing.

[0063] In a corresponding embodiment of the aircraft according to the invention, the rear wing thus comprises a dihedral angle (ε), at least in the outer part of the rear wing, and preferably increasing towards the outer part of the rear wing (i.e. towards said connecting beam), preferably, 0° < ε < 5°

[0064] In aeronautics, the "dihedral angle" is the angle between the left and right wings (or tail surfaces) of an aircraft. In particular, the dihedral angle is the angle upwards from the horizontal plane of the wing or the horizontal tail of a (fixed-wing) aircraft. The "anhedral angle" is the name for a negative dihedral angle, i.e. the anhedral angle is the angle downwards from the horizontal plane of the wing or the horizontal tail of a (fixed-wing) aircraft.

[0065] In a corresponding embodiment of the aircraft according to the invention, the aircraft further comprises at least one forward propulsion unit, which preferably (but not limited to) has at least one propeller and at least one motor driving the propeller, preferably an electric motor, and more preferably, the forward propulsion unit is located at the rear of the aircraft. Such a forward propulsion unit facilitates forward flight.

[0066] In a preferred embodiment of the aircraft according to the invention, there is at least one forward propulsion unit on each side of the fuselage, preferably, the forward propulsion unit is located below the front wing and the rear wing, and most preferably, the forward propulsion unit is located in front of the rear wing. This has proven to be very beneficial for achieving a high forward flight speed.

[0067] In a further advantageous embodiment of the aircraft according to the invention, when viewed along the longitudinal axis of the aircraft, the front wing and the rear wing together form a closed-loop shape. This increases mechanical stability and reduces negative aerodynamic effects. Description of the Drawings

[0068] Further features and advantages of the invention can be derived from the following description of exemplary embodiments with reference to the drawings.

[0069] Figure 1 A plan view of an aircraft according to the invention is shown;

[0070] Figure 2 A side view of an aircraft according to the invention is shown;

[0071] Figure 3 A front view of an aircraft according to the invention is shown;

[0072] Figure 4 Shows Figure 1 a simplified version of

[0073] Figure 5 Another front view of an aircraft according to the invention is shown; and

[0074] Figure 6Shows another plan view of an aircraft according to the present invention. Detailed implementation

[0075] Figure 1 Shows a vertical takeoff and landing aircraft 1 having a fuselage 2 or a fuselage body for transporting passengers and / or loads. x, y, and z represent the aircraft axes (longitudinal, lateral, and vertical respectively, as described above). The front wing 3 is attached to the fuselage 2. In addition, the rear wing 4 is attached to the fuselage 2, and the rear wing 4 is behind the front wing 3 in the forward flight direction (arrow FF) of the aircraft 1. The right connecting beam 5a and the left connecting beam 5b are structurally connected to the front wing 3 and the rear wing 4, and the connecting beams 5a, 5b are laterally spaced apart from the fuselage 2. The aircraft 1 includes six lifting units M1-M6, with three lifting units provided on each of the connecting beams 5a, 5b. Each lifting unit M1-M6 includes at least one propeller 6b and at least one motor 6a driving the propeller 6b. The motor is preferably an electric motor, and the lifting units are arranged such that their respective propeller axes are in a substantially vertical direction (z-axis). This is Figure 1 Only one of the lifting units M1-M6 (M6) is clearly shown in

[0076] The lifting units M1-M6 are placed symmetrically with respect to the longitudinal (x) and lateral (y) axes of the aircraft 1: When compared with the lifting unit M3, the lifting units M1 and M5 are placed forward and backward respectively (in the x direction) at equal distances. The lifting units M2, M4, and M6 are also placed in the same way. In addition, the three lifting units on each side of the aircraft 1 are placed at equal lateral distances (in the y direction) from the longitudinal axis of the aircraft.

[0077] The reference numerals P1-P6 denote the rotor planes, which can be defined as the respective regions swept by the rotating propellers 6b of the lifting units M1-M6. The wing attachments of the connecting beams 5a, 5b are placed between the rotor planes P1-P6 and outside the rotor planes P1-P6. The wing attachment for the front wing 3 is denoted as A3, and the wing attachment for the rear wing 4 is denoted as A4.

[0078] Both the front wing 3 and the rear wing 4 (at least a part of them, for example, the outer part of the rear wing 4 relative to the fuselage 2) define sweep angles γ, β (forward sweep) less than 90°. The rear wing 4 has a rear sweep (angle δ) in the region near the fuselage 2 (V-shaped tail), which helps to reduce the length of the fuselage 2 while maintaining a sufficient rear surface wing area S 后 (See below). The preferred values of these angles are 45° < γ < 135°, 30° < β < 90°, and 70° < δ < 170°. Preferably, γ = 75°, preferably, β = 65°, and preferably, δ = 110°.

[0079] The reference signs F1 and F2 denote forward propulsion units (thrusters) which are respectively attached to the left and right sides of the fuselage 2 and are close to the rear of the aircraft 1 (in front of the rear wing 4). The thrusters may include at least one propeller and at least one motor (not shown) for driving the propeller.

[0080] The specific physical points of the aircraft 1 defined previously are denoted as CoG (center of gravity), CoA (center of drive), CA (aerodynamic center), and NP (neutral point). The aerodynamic centers (CA 前 ) of the front wing 3 and the rear wing 4 (CA 后 ) are also denoted, and are shown on that part of each wing which extends between the fuselage 2 and the right connecting beam 5a. The same applies to that part of each wing which extends between the fuselage 2 and the left connecting beam 5b ( Figure 1 , not shown). As can be seen from Figure 1 , the center of drive CoA of the aircraft 1 is geometrically in the same longitudinal position (x CoA ) as (at least approximately) the center of gravity CoG of the aircraft. Due to the forward sweep of the front wing 3 and the rear wing 4, the neutral point NP is well behind the center of drive CoA and the center of gravity CoG.

[0081] The wing area S 前 of the front wing 3 is less than the wing area S 后 of the rear wing 4. In the present example, 60% < S 前 / S 后 < 70%. In summary, the rear wing 4 includes the wing surfaces of the V-shaped tail (half wing 4b) and the wing surfaces of the conventional lifting wing (half wing 4a), which will become apparent from the following Figures 3 to 5 .

[0082] Figure 2 A side view of the aircraft 1 is shown in Figure 1 . In all the figures, the same elements are denoted by the same reference signs.

[0083] The connecting beams (only the beam 5b is visible in Figure 2 ) are placed higher in the z direction than or above the fuselage 2 (see Figure 1 ).

[0084] M1 + 2 denotes the lifting units M1 and M2, one of which is behind the other in the line of sight. The same applies respectively to the lifting units M3 and M4 and the lifting units M5 and M6.

[0085] The reference sign D denotes the propeller downwash (only one of the lifting units M1 - M6 (M5 + 6) is shown). The above-mentioned wing attachments (see Figure 1 ) are located outside the downwash area.

[0086] Figure 3 A front view of the aircraft 1 is shown. It can be seen that in a plane perpendicular to the longitudinal axis (x) of the aircraft 1, the rear wing 4 extends laterally from the fuselage 2 such that the angle θ formed by the respective half-wings 4a, 4b of the rear wing 4 is less than 180°, preferably approximately 90° to 100°. The inner half-wing 4b forms the V-tail fin surface, while the outer half-wing 4a represents a conventional elevator wing surface.

[0087] The front wing 3 includes a dihedral angle (λ), which in combination with the above-mentioned sweep angle (γ < 90°), has a positive effect on the lateral / directional stability of the aircraft 1. This angle λ can increase towards the respective tips of the front wing 3. In addition, the rear wing 4 includes anhedral angle (ε) in order to generate a high vertical offset VO in the region of the higher downwash flow 7 of the front wing 3 (see Figure 2 ). This especially occurs in the inner part (spanwise direction) of the front wing 3. This angle ε can also increase towards the respective tips of the rear wing 4. Thus, when viewed along the longitudinal axis (x) of the aircraft 1, the front wing 3 and the rear wing 4 together form a closed-loop shape.

[0088] Figure 4 is Figure 1 a slightly simplified version for illustrating the wing area S of the front wing 3 前 is smaller than the wing area S of the rear wing 4 后 . However, due to the V-tail configuration, the vertical lift generated by the rear wing surface is smaller. The rear wing 4 includes a V-tail fin surface (half-wing 4b) and a conventional elevator wing surface (half-wing 4a), which are represented by different shadings.

[0089] This is further shown in Figure 5 which is another front view of the aircraft 1 (for more details see Figure 5 Figure 3 Figure 3 ).

[0090] Finally, as shown in Figure 6 , the V-tail sweep angle δ, δ > 90°, also promotes the improvement of load transfer (the dashed line indicated by reference numeral 8), because the "V" configuration transfers the load from the connecting beam 5b through the wing 4a and the V-tail 4b to the fuselage 2.

Claims

1. A vertical takeoff and landing aircraft, characterized in that, Comprising: A fuselage (2) for transporting passengers and / or a load; A front wing (3) attached to the fuselage (2); A rear wing (4) attached to the fuselage (2), with the rear wing (4) being behind the front wing (3) in the forward flight direction (FF); A right connecting beam (5a) and a left connecting beam (5b), the connecting beams (5a, 5b) being structurally connected to the front wing (3) and the rear wing (4), and the connecting beams (5a, 5b) being spaced apart from the fuselage (2); And At least two lifting units (M1 - M6) on each of the connecting beams (5a, 5b), each lifting unit (M1 - M6) including at least one propeller (6b) and at least one motor (6a) driving the propeller (6b), the lifting units (M1 - M6) being arranged such that their respective propeller axes are in a substantially vertical direction (z); Wherein at least part of the front wing (3) has a sweep angle γ, 45° ≤ γ ≤ 135°, the rear wing (4) includes outer rear wing portions (4a) on each side of the fuselage (2), and the outer rear wing portions (4a) as a whole have a forward sweep with a sweep angle β ≥ 30°; The rear wing (4) includes inner rear wing portions (4b) from the fuselage (2) to the outer rear wing portions (4a); Wherein the inner rear wing portions (4b) are formed in the form of a V - tail with a sweep angle (δ) greater than 90°; And, in a plane perpendicular to the longitudinal axis (x) of the aircraft (1), the rear wing (4) extends laterally from the fuselage (2) such that the corresponding inner rear wing portions (4b) forming the lateral stability unit of the aircraft have an angle (θ) satisfying 90° < θ < 100°; Wherein the front wing (3) includes a dihedral angle (λ), and the outer rear wing portions (4a) of the rear wing (4) include anhedral angles (ε) towards the connecting beams (5a, 5b).

2. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The wing area (S 前 ) of the front wing (3) is smaller than the wing area (S 后 ) of the rear wing (4).

3. The vertical takeoff and landing aircraft according to claim 1, characterized in that, Each of the connecting beams (5a, 5b) includes three of the lifting units (M1 - M6).

4. The vertical takeoff and landing aircraft according to claim 1, wherein, The lifting units (M1 - M6) are symmetrically placed with respect to the longitudinal plane (x - z) and the lateral plane (y - z) of the aircraft (1).

5. The vertical takeoff and landing aircraft according to claim 1, wherein Attachments (A3, A4) of the front wing (3) and the rear wing (4) respectively connected to the connecting beams (5a, 5b) are placed between the respective rotor planes (P1 - P6) swept by the propellers (6b).

6. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The rear wing (4) is higher than the front wing (3) in the vertical position (z).

7. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The front wing (3) includes the dihedral angle (λ) at the respective tips of the front wing (3).

8. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The anhedral angle (ε) is included on the outer part of the rear wing (4).

9. The vertical takeoff and landing aircraft according to claim 1, wherein The longitudinal aircraft stability defined by the quantity Cm / α is negative, where Cm is the pitching moment coefficient and α is the angle of attack.

10. The vertical takeoff and landing aircraft according to claim 1, characterized in that, The center of actuation (CoA) is geometrically in the same longitudinal position (x CoA ) as the center of gravity (CoG) of the aircraft, and the neutral point (NP) is located behind the center of actuation (CoA) and the center of gravity (CoG).

11. The vertical takeoff and landing aircraft according to claim 1, wherein The vertical take-off and landing aircraft further includes at least one forward propulsion unit (F1, F2), the forward propulsion unit (F1, F2) having at least one propeller and at least one motor for driving the propeller, and the forward propulsion unit (F1, F2) is located at the rear of the aircraft (1).

12. The vertical takeoff and landing aircraft according to claim 11, wherein, There is at least one of the forward propulsion units (F1, F2) on each side of the fuselage (2), and the forward propulsion unit (F1, F2) is located below the front wing (3) and the rear wing (4).

13. The vertical takeoff and landing aircraft according to claim 1, characterized in that, When viewed along the longitudinal axis (x) of the aircraft (1), the front wing (3) and the rear wing (4) together form a closed-loop shape.

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