Aircraft with folding mechanism

Abstract

An aircraft with a folding mechanism is proposed, comprising: a fuselage; a payload and / or landing gear, optionally connected to the underside of the fuselage; at least two longitudinal beams connected to the upper and / or sides of the fuselage, the at least two longitudinal beams extending parallel to each other and parallel to a first axis of the aircraft, each of the longitudinal beams connecting to a plurality of lifting units; at least one crossbeam connected to the upper side of the fuselage, the at least one crossbeam extending parallel to a second axis of the aircraft and perpendicular to the longitudinal beams, the crossbeam connecting to the plurality of lifting units; wherein the longitudinal beams are rotatably connected to the fuselage via at least one corresponding first pivot joint, the first pivot joint being designed for pivoting the longitudinal beams about the corresponding first pivot axis to a pivot position; and the crossbeams are rotatably connected to the fuselage via at least one second pivot joint, the second pivot joint being used for pivoting the crossbeams about the second pivot axis to a pivot position.

Description

Technical Field

[0001] This invention relates to an aircraft with a folding mechanism, particularly an electric vertical take-off and landing (eVTOL) aircraft, i.e., an aircraft with multiple electric lift / propulsion units and vertical take-off and landing capabilities. Such an aircraft is also commonly referred to as a "multi-rotor aircraft". Background Technology

[0002] DE 102012202698A1 discloses an aircraft of the above type, which has a boom with a lifting unit arranged on the boom, the lifting unit being in the form of an electric motor with a propeller (rotor). The boom is connected to a central hub and can be detached or folded for storage or transport.

[0003] Figure 1 A perspective view (top view) and a plan view (bottom view) of an aircraft of the aforementioned type in the prior art are shown. Specifically, Figure 1 An aircraft 1, such as the multi-rotor electric vertical takeoff and landing aircraft presented by the applicant, is shown. The aircraft 1 includes a fuselage 2, which, among other things, houses an onboard flight controller 2a for controlling multiple, for example 18, electrically propelled or lifting units 3. Each propulsion unit includes at least one motor 3a and a rotor 3b (preferably integral) with corresponding rotor blades. For clarity, Figure 1 Only one propulsion (lifting) unit 3 and its components are explicitly shown. Propulsion unit 3 is located on the branched Y-shaped arm 3c. Reference numeral 4 indicates an exemplary sensor unit used to measure the current state of aircraft 1 and / or propulsion unit 3. Reference numeral 5 indicates a connecting element used to interconnect adjacent arms 3c. Arm 3c is connected to the central hub 6. As shown, sensor unit 4 and propulsion unit 3 are operatively communicatively connected to flight controller 2a. Reference numerals x, y, and z indicate the main control axes of the aircraft: roll (longitudinal), pitch (lateral), and yaw (vertical). The corresponding moments are represented by L, M, and N.

[0004] For example, existing folding and rotating mechanisms known from prior art document DE 202006017959U1 require heavy additional equipment to be carried during flight, thus limiting flight range and / or effective payload.

[0005] Lightweight and easy-to-operate folding mechanisms (ideally operable by only one person) would be advantageous for heavy-duty drone applications, such as in agriculture, construction sites, or logistics, where aircraft / drones are typically transported to their respective sites, assembled, used, disassembled, and then stored on trailers. However, folding and rotating mechanisms known to be used in aircraft require extreme care to ensure they are used in the correct orientation. Furthermore, these mechanisms often include fragile longitudinal beams, necessitating sophisticated ground clamps to secure them. Additionally, complex fixing devices may be required to anchor the beams to the floor.

[0006] This makes moving the folded drone onto a trailer for transport difficult or cumbersome. The assembled drone needs to be landed on the trailer or moved before it can be disassembled.

[0007] From the perspective of load path, in the current known flight configuration of electric vertical takeoff and landing aircraft, the main loads transferred from side to side and from front to rear are carried out through the central hub, thus increasing the connection load. This high connection load leads to wear on all the connecting elements involved (such as bolts, bushings, etc.). Summary of the Invention

[0008] The purpose of this invention is to overcome the above-mentioned disadvantages and to provide an aircraft with a folding mechanism that increases durability and stability while reducing weight.

[0009] This objective is achieved by the aircraft provided by the main embodiment of the invention. Advantageous further embodiments are defined in the dependent embodiments.

[0010] According to the present invention, an aircraft with a folding mechanism, particularly an electric vertical takeoff and landing (eVTOL) aircraft, comprises: a fuselage, which may include a payload bay; optionally, a payload bay and / or landing gear connected to the fuselage, preferably, the payload bay and / or landing gear located below the fuselage; at least two longitudinal beams connected to the fuselage, preferably located above and / or to the side of the fuselage, preferably, the at least two longitudinal beams extending parallel to each other and parallel to a first axis of the aircraft, each of the longitudinal beams being connected to a plurality of lifting units; and at least one crossbeam connected to the fuselage, preferably located above the fuselage, and The at least one crossbeam extends parallel to the second axis of the aircraft and preferably at a right angle to the longitudinal beams (if the longitudinal beams are parallel to each other and parallel to the first axis of the aircraft), the crossbeams being connected to a plurality of lifting units; wherein the longitudinal beams are rotatably connected to the fuselage via at least one corresponding first pivot joint, the first pivot joint being designed to pivot the longitudinal beams about a corresponding first pivot axis to a pivot position; and wherein the crossbeams are rotatably connected to the fuselage, preferably rotatably connected to the fuselage via at least one second pivot joint, the second pivot joint being used to pivot the crossbeams about a second pivot axis to a pivot position.

[0011] Preferably, the first axis can be the longitudinal axis of the aircraft, and the second axis can be the transverse axis of the aircraft. However, the axes can also be interchanged.

[0012] The first pivot axis can be parallel to the longitudinal axis of the aircraft, while the second pivot axis can be parallel to the vertical axis of the aircraft; this is a preferred configuration.

[0013] The basic design principles disclosed in this specification can be followed without using a beam, but the use of such a beam is preferred for stability. In any case, designing an aircraft without such a beam would also be a feasible modification of this disclosure.

[0014] Given that existing technology includes a central hub structure combining a sliding mechanism and a rotating mechanism, wherein the arms or rotor beams are mounted to the central hub, the new design has at least two preferably identical (i.e., mirrored or flipped) longitudinal rotor beams and a crossbeam, but without a central hub. The crossbeam extends parallel to the transverse axis of the aircraft and perpendicular to the longitudinal beams. The longitudinal beams and crossbeam are also referred to as "rotor beams". As previously mentioned, the rotor beams can be interconnected using ring connectors (connecting elements) to increase stability.

[0015] The load paths from left to right and from front to back are achieved through the longitudinal beams and / or the transverse beams, and the interface loads of the fuselage are generated by the payload (optional) and the weight of the fuselage (combination). There is no need to transfer any flight payload to the other side of the rotorcraft via the fuselage.

[0016] The possible features and advantages of the proposed structural concept for a foldable structure for (large) multi-rotor aircraft may include:

[0017] • A ring connector (connecting element) that folds inward and can be locked to other beams.

[0018] • The longitudinal beams fold downwards;

[0019] • The crossbeam rotates forward and locks in that position;

[0020] • The cable does not need to be disconnected during transportation;

[0021] • No major components need to be disassembled for transport;

[0022] • All connectors can be installed using quick-release fasteners without any tools;

[0023] • The length, width, and height of the folding system (aircraft) allow for container transport (e.g., with 40-foot containers) and also allow for traditional road transport.

[0024] This invention is not limited to aircraft with any particular number of lift units, although examples include multi-rotor aircraft with 18 motors (N=18), hexacopter aircraft (N=6), and octocopter aircraft (N=8). Other configurations are also possible.

[0025] As used in this specification, the term "longitudinal beam" generally refers to the beam running from the front to the rear of a multirotor aircraft (preferably parallel to the x-axis, see [reference]). Figure 1 The beam on which the motor is mounted. This beam is preferably folded downwards (pivot). "Crossbeam" refers to a beam that typically runs from left to right (lateral). Figure 1 The crossbeam that mounts the motor (on the y-axis). It is best to rotate it 90° during folding. "Ring connectors" or "connecting elements" refer to the structural interconnecting beams between the rotor beams. They can be used in larger multirotor aircraft to increase stability, but they are not mandatory.

[0026] In one embodiment of the aircraft according to the invention, the lifting unit includes at least one corresponding rotor, the plurality of rotors being arranged in a common rotor plane, serving at least respectively for the longitudinal beam and / or the crossbeam. The rotor plane may be different between the longitudinal beam on one side and the crossbeam on the other side. A single rotor may be tilted relative to the plane.

[0027] In a specific embodiment of the aircraft according to the invention, the crossbeam is along the vertical axis of the aircraft ( Figure 1 The z-axis (in the equation) is arranged above the longitudinal beam.

[0028] In another embodiment of the aircraft according to the invention, the longitudinal beam has multiple branches, and the lifting units are located at corresponding bifurcation points and / or endpoints of the branches. Preferably, 2x3 lifting units are arranged on a designated beam, which may have a (double)Y-shaped configuration, and the lifting units arranged on the common branch are arranged in at least one triangular configuration.

[0029] In another embodiment of the aircraft according to the invention, the crossbeam has multiple branches, and the lifting units are located at corresponding bifurcation points and / or endpoints of the branches, preferably arranged in at least one triangular configuration. Alternatively, the crossbeam may have a (double) Y-shaped configuration, and the lifting units disposed on the common branches are arranged in at least one triangular configuration.

[0030] In another embodiment of the aircraft according to the invention, any rotor included in the lifting unit can be connected to transport the aircraft in rotational positions relative to the longitudinal beam and the crossbeam, respectively. This can be achieved through a strap system, dedicated protective accessories, dust covers, etc.

[0031] In another advantageous embodiment of the aircraft according to the invention, the rotors located on the longitudinal beam can be connected (for transport) such that their blades are oriented parallel to the longitudinal beam, while the rotors located on the crossbeam can be connected (for transport) such that their blades are parallel to the crossbeam and oriented laterally to the longitudinal beam. This facilitates folding operations while protecting the rotor blades. Furthermore, this ensures that the rotors do not hit the ground during folding. In this context, "connectable" means that the rotors can be connected or secured to the crossbeam so that they do not move during folding and / or transport.

[0032] In yet another embodiment of the invention, the aircraft includes a plurality of connecting elements for interconnecting one of the plurality of longitudinal beams with the crossbeam, or for interconnecting one of the plurality of longitudinal beams with another of the plurality of longitudinal beams, via each connecting element. These connecting elements (“ring connectors”) have already been mentioned previously. They provide enhanced structural stability, especially in larger multirotor aircraft (e.g., N≥8).

[0033] In a particularly advantageous further embodiment of the aircraft according to the invention, a designated connecting element is rotatably connected to one of the plurality of longitudinal beams or to the crossbeam via a third pivot joint, the third pivot joint being designed for rotating the connecting element to abut against one of the plurality of longitudinal beams or the crossbeam. Thus, when folding the aircraft, it can be easily folded as a first step.

[0034] In a further embodiment, the aircraft may include a quick-release mechanism for securing the connecting element to the crossbeam. Preferably, the quick-release mechanism includes a pin that passes through alignable through-holes in both the connecting element and the crossbeam. More preferably, the pin passes through alignable through-holes in both the connecting element and a corresponding bracket connected to the crossbeam. This facilitates the folding operation of the connecting element. The bracket can be used to accommodate the height difference between the crossbeam and the longitudinal beam.

[0035] Specifically, the rotary interface included in the quick-release mechanism can be secured with a bolt (or pin) fastened with a nut or similar material, and once the quick-release interface is unlocked, the rotary interface is considered to be able to rotate around the z-axis (or an axis parallel to the z-axis, see [link]). Figure 1 It can be released by free rotation.

[0036] While the quick-release interface can be implemented with a bolt secured by a nut, quick-release solutions including, for example, quarter-rotary bolts or pin / cotter pin combinations may be preferred.

[0037] After the folded ring connectors are attached, they can be advantageously secured to the rotor beam for transport, and the aircraft may include a corresponding securing mechanism.

[0038] Since the vertical position of the crossbeam can differ from that of the longitudinal beam, brackets (mentioned earlier) can be used on the crossbeam to allow the use of the same ring connector.

[0039] During folding, all quick-release pins (or equivalents) must be opened to separate the longitudinal beams, which can then be folded downwards. The quick-release mechanism of the longitudinal beams can be combined with the unlocking system of the crossbeams to reduce weight.

[0040] Ideally, all substructures used to secure the longitudinal beams should transfer the corresponding load directly to the fuselage bulkheads, which are directly connected to the (optional) landing gear and any cargo mounting hardware (e.g., supports for attaching the payload). This results in an optimized load path.

[0041] In another embodiment of the aircraft according to the invention, the aircraft includes an electrical connector, such as a cable, located between the fuselage and the elevator unit, wherein the electrical connector has sufficient overhang near the first pivot joint and / or the second pivot joint for pivoting the longitudinal beam and the crossbeam respectively, without removing the electrical connector during folding. This also applies to any electrical connector used with sensors or other electrical equipment deployed on the rotor beam. The cable routing is such that the connection does not need to be broken during folding.

[0042] In another embodiment of the aircraft according to the invention, the longitudinal beam surrounds the fuselage and optionally connects to the landing gear and / or payload at its pivot position. While this may help protect the fuselage and / or payload / landing, it also provides an advantageous configuration with minimal space expansion.

[0043] If the aircraft includes a payload bay as described above, the payload can be located therein during transport.

[0044] The downward folding mechanism can be supported by a gas spring or similar device to allow for single-person operation. The gas spring can be part of the aircraft, or (preferably) designed as a separate assembly so that it does not add any weight to the aircraft during flight.

[0045] In a corresponding embodiment of the aircraft according to the invention, the aircraft includes an elastic mechanism, such as a corresponding gas spring, designed to support the pivoting movement of the longitudinal beam, preferably in the form of a detachable mechanism. In particular, this can further facilitate the folding operation if it is performed by a single person.

[0046] In another embodiment of the aircraft according to the invention, the aircraft includes a detachable beam clamp that provides the second pivot joint. Because the clamp is detachable, it does not add weight to the aircraft during flight.

[0047] In another embodiment of the aircraft according to the invention, the aircraft includes a fixing device for securing the crossbeam to the longitudinal beam at the pivot position. This provides additional safety during transport.

[0048] As previously described, the aircraft according to the invention may include N=6, N=8, or N=18 (but not limited to) lift units, each lift unit having at least one rotor. Generally, the aircraft may include any even number of lift units, N=2k, where k∈N, preferably N=6, N=8, or N=18, wherein: when N=6, two lift units are provided at opposite ends of the crossbeam, and two lift units are provided at opposite ends of each of the two longitudinal beams; when N=8, four lift units are provided on the crossbeam, preferably at their respective bifurcation points and / or the endpoints of the branches, and two lift units are provided at opposite ends of each of the two longitudinal beams; when N=18, six lift units are provided on the crossbeam, preferably at their respective bifurcation points and / or the endpoints of the branches, and six lift units are provided on each of the two longitudinal beams, preferably at their respective bifurcation points and / or the endpoints of the branches.

[0049] To complete the folding operation, the crossbeam, which has been separated from the longitudinal beam, revolves around the vertical z-axis (see...). Figure 1 Rotate forward 90°. The corresponding pivot point (or pivot joint) can be a rotation point fixedly mounted on the fuselage, or, preferably, can be designed to install only clamps for rigging / unrigging so as not to increase the weight of the aircraft during flight.

[0050] After rotating the crossbeam, its corners are preferably secured to the longitudinal beam to stabilize and reinforce the structure during transport. For this purpose, existing mounting locations for the ring connector can be advantageously utilized.

[0051] The beam can be rotated manually without using a dedicated pivot point (or pivot joint) at the center of the aircraft. While this would require two people, it could be advantageous because it eliminates the need for an additional weight-bearing rotating pivot system.

[0052] Other features and advantages of the invention will become apparent from the following description of preferred embodiments with reference to the accompanying drawings. Attached Figure Description

[0053] Figure 1 This illustrates a conventional aircraft configuration;

[0054] Figure 2 A schematic diagram of an aircraft according to the present invention is shown;

[0055] Figure 3 Different aircraft configurations according to the present invention are shown;

[0056] Figure 4 It shows Figure 2 The location of quick-release and rotary interfaces in aircraft;

[0057] Figure 5 The folding is shown Figure 2 The first step of the ring connector for the aircraft;

[0058] Figure 6 An embodiment of a quick-release mechanism for a ring-shaped connecting element is shown;

[0059] Figure 7 It shows Figure 2 The second step in the folding process of the aircraft;

[0060] Figure 8 It shows a structure with folded longitudinal beams. Figure 2 The aircraft (the third step in the aircraft folding process);

[0061] Figure 9 Details of the aircraft's structure are shown;

[0062] Figure 10 It shows Figure 2 The fourth step in the folding process of an aircraft;

[0063] Figure 11 The details of the aircraft's cable wiring are shown; and

[0064] Figure 12 It shows Figure 2 The final folded state of the aircraft. Detailed Implementation

[0065] Figure 1 It has been further described; it shows the existing technology of aircraft 1.

[0066] In the following figures, the same reference numerals denote the same elements or elements that provide at least similar functions. In the case of the same element, not all elements are provided with reference numerals to increase comprehensibility.

[0067] Figure 2 The diagram schematically illustrates a novel aircraft design without a central hub. The aircraft 1 includes a fuselage 2, beneath which payload and / or landing gear can be connected (see [reference]). Figure 12 As shown in the figure, the aircraft 1 also has at least two longitudinal beams 7.1 and 7.2 connected to the fuselage 2, preferably above the fuselage 2 and on the side of the fuselage 2. The longitudinal beams 7.1 and 7.2 are parallel to each other and parallel to the longitudinal axis of the aircraft (in...). Figure 2(represented by x in the figure) extends. Each of the longitudinal beams 7.1 and 7.2 connects to multiple lifting units 3, only one of which is indicated by reference numerals 3 and 3b (where 3b represents the corresponding rotor by the corresponding propeller sweep area). Therefore, each lifting unit in Figure 2 The dashed circle indicates the rotor. Any rotor 3b located on the longitudinal beams 7.1 and 7.2 can be arranged in a common rotor plane.

[0068] At least one crossbeam 8 is also connected to the fuselage 2, preferably above the fuselage 2 and above the longitudinal beams 7.1, 7.2, and it is parallel to the transverse axis of the aircraft 1 (in Figure 2 (represented by y in the figure) extends perpendicularly to the longitudinal beams 7.1 and 7.2. As shown in the figure, the crossbeam 8 connects multiple lifting units. Any rotor 3b located on the crossbeam 8 can be arranged in a common rotor plane, which can (but does not have to) be the same as the rotors located on the longitudinal beams 7.1 and 7.2. The longitudinal beams 7.1 and 7.2 are connected by at least one corresponding first pivot joint (see Figure 8 The first pivot joint is rotatably connected to the fuselage 2, and is designed to pivot the longitudinal beams 7.1, 7.2 about a corresponding first pivot axis parallel to the longitudinal axis (x). Furthermore, the crossbeam 8 is rotatably connected to the fuselage 2, preferably via at least one second pivot joint (see [link to documentation]). Figure 10 The second pivot joint is connected to the fuselage 2 and is used to pivot the beam 8 about the second pivot axis to the pivot position, the second pivot axis being parallel to the vertical axis (z) of the aircraft 1.

[0069] The longitudinal beams 7.1 and 7.2 and the transverse beam 8 all have branched structures, and they all face their respective free ends (i.e., away from fuselage 2) in a "Y" shape. Furthermore, there are additional branches connecting to the upper Y-shaped legs, thus forming a triangular beam configuration at the ends of the longitudinal beams 7.1, 7.2, and transverse beam 8. Figure 1 Similarly, the various branches or beams of the Y-shaped beam section 3c are indicated by the reference numeral 3c', which includes those beams that interconnect the upper ends of the Y-shaped legs. However, Figure 2 There is no central hub, and each beam, namely the longitudinal beams 7.1, 7.2 and the transverse beam 8, includes two Y-shaped portions 3c at its end. The lifting unit 3 is located at the corner (apex) of the triangular beam configuration. As in the prior art, elements 5 (ring connectors or connecting elements) interconnect adjacent Y-shaped beam portions 3c to form an integral ring configuration to increase stability.

[0070] Although in the illustrated embodiment, the beam 8 is about its longitudinal axis (from the free end to the free end, i.e. along...) Figure 2The y-axis in the image is perfectly symmetrical from left to right, but as... Figure 2 As shown, longitudinal beams 7.1 and 7.2 have an asymmetrical configuration in this respect. In fact, they are designed as identical components and arranged symmetrically with respect to the central transverse axis (parallel to beam 8). They are further arranged in a mirror-image manner (relative to the longitudinal axis of aircraft 1; see...) Figure 2 (x-axis in the diagram).

[0071] Preferably, the pivot axes for folding the longitudinal beams 7.1 and 7.2 downwards are arranged parallel to the longest straight-line extension direction of the longitudinal beams 7.1 and 7.2. These pivot axes are connected by... Figure 2 The dotted line PA is used to depict this. This is advantageous because it allows the longitudinal beams 7.1 and 7.2 to be folded downwards without any rotor 3b extending upwards beyond the plane defined by the pivot axis PA. This will become clearer below (see, for example, [link]). Figure 8 and Figure 12 ).

[0072] Figure 3 The basic design principles of the aircraft 1, which has six, eight, and eighteen lift units 3 (from left to right), are explained.

[0073] Figure 4 It shows Figure 2 The positions of the aircraft 1, quick-release interface 5a, and rotary interface (or pivot connector) 5b connected to the annular connector 5. When the corresponding quick-release interface 5a is disengaged, the annular connector 5 can rotate around the rotary interface 5b, as shown below. Figure 5 As indicated by arrow R in the diagram. The ring connector 5 is therefore placed on the Y-shaped beam section or branch 3c', where, as described above, they can be connected for transport.

[0074] Figure 6 Details regarding possible embodiments of the quick-release interface 5a are provided. The crossbeam 8 is preferably provided with a bracket 8a that can compensate for different vertical positions (heights) of the crossbeam 8 and the longitudinal beam (not shown). The bracket 8a and the annular connector 5 have alignable through holes 8aa, 8ab, 5c through which cotter pins or pins (bolts) 9, secured by nuts 10, pass.

[0075] Figure 7 The diagram illustrates a second step for pre-positioning the rotors 3b during the folding of the aircraft 1. For those rotors 3b located on the longitudinal beams 7.1 and 7.2, these rotors 3b are oriented (preferably, and fixed) such that their respective blade axes point in the longitudinal direction x, while for those rotors 3b located on the transverse beam 8, these rotors 3b are oriented such that they point in the transverse direction y. Alternatively, the rotors can be disassembled prior to the second step.

[0076] Then, according to Figure 8 In the third step, longitudinal beams 7.1 and 7.2 are folded downwards, as indicated by arrow R'. Reference numeral 7a indicates a (quick) separation mechanism for longitudinal beam 7.1, which must be operated before rotation R' can be performed. The same applies to longitudinal beam 7.2. Separation mechanism 7a can be designed as follows: Figure 6 The separation mechanism 5a is shown. For rotation, the longitudinal beams 7.1 and 7.2 are fixed to their respective rotation points or pivot joints PR (shown only on one side), which are connected to the fuselage 2 via the support structure 7b. Therefore, the corresponding pivot axes are parallel to the longitudinal axis of the aircraft (see [reference]). Figure 2 (The axis PA in the figure). The reference numerals 7.1' and 7.2' indicate the longitudinal beam after rotation.

[0077] Figure 8 The other separation mechanism 5a shown can be the one connected in front. Figure 4 and Figure 6 Those mentioned (used for separating the ring connector 5). Some of them, namely those located on the longitudinal beams 7.1 and 7.2, can be reused to secure the folded longitudinal beams 7.1' and 7.2' to the landing gear 2'' located below the fuselage 2 (and the optional payload 2'). For this purpose, the landing gear 2'' can have a suitable fixing structure (not shown). The landing gear 2'' (and the aforementioned separation mechanism 5a) defines the maximum width MW of the folded aircraft 1 for transport purposes.

[0078] The aircraft 1 may include an elastic mechanism, such as a corresponding gas spring, designed to support the pivoting movement of the longitudinal beams 7.1, 7.2, preferably in the form of a detachable mechanism. This is not shown in the figures.

[0079] Figure 9 Further structural details of the aircraft 1 are shown. Reference numeral 8b indicates a beam support for securely connecting the beam 8 to the fuselage bulkhead 2a. Landing gear support 2a'' for connecting the landing gear 2'' to the fuselage 2 (or bulkhead 2a) and a payload mounting bracket 2a'' for connecting the payload 2'' to the fuselage 2 (or bulkhead 2a or landing gear 2'') are also shown. Reference numeral 100 indicates the ground.

[0080] Figure 10 The next (fourth) step in the folding process of aircraft 1 is shown, similar to... Figure 7 The diagram is in the image, but the longitudinal beam is already in its downward rotation position (see...). Figure 8 The crossbeam 8 is rotated along the longitudinal x direction according to arrow R'' to its new position, denoted as 8'. According to the crossbeam support 8b ( Figure 9The property of R'' can be achieved by separating the crossbeam 8 and then rotating it about a pivot point PR', which can be a fixed pivot joint. However, embodiments without such a joint are possible and have been further described above. In particular, the aircraft 1 may include a detachable crossbeam clamp (not shown) that temporarily provides the pivot joint for rotating R''. The rotation R'' is about a pivot axis parallel to the vertical axis (z) of the aircraft.

[0081] At reference numeral 5a in the attached drawing, the separation mechanism on the crossbeam 8 can be reused to provide interconnection between the rotating crossbeam 8' and the (rotating) longitudinal beams 7.1', 7.2' for transport purposes. However, a dedicated additional mechanism (not shown) may also be provided.

[0082] Figure 11 Includes details from the final step, which shows cable 11 passing along (or within) the crossbeam 8 and then through mounting piece 8a (see...). Figure 9 The cable enters the fuselage (not shown). The cable can provide power to the rotor and / or signal connections to sensors or other devices (see [link to documentation]). Figure 1 Cable 11 has sufficient (slack) length to avoid breakage and / or disconnection when the crossbeam 8 is rotated to its rotation position 8'. Reference numeral 11' indicates the cable after the crossbeam has been rotated.

[0083] Figure 12 The final folded configuration of Aircraft 1 is shown, as explained in detail above. On the left, Aircraft 1 is shown in a top view. On the right, Aircraft 1 is shown in a front view. Advantageously, the maximum width and length allow the folded Aircraft 1 to fit into a standard 40-foot container for transport.

Claims

1. An aircraft having a folding mechanism, the aircraft (1) comprising: fuselage (2); At least two longitudinal beams (7.1, 7.2) are connected to the fuselage (2), the at least two longitudinal beams (7.1, 7.2) are parallel to each other and extend parallel to the longitudinal axis (x) of the aircraft (1), and each of the longitudinal beams (7.1, 7.2) is connected to a plurality of lifting units (3). At least one crossbeam (8) is connected to the fuselage (2), the at least one crossbeam (8) is parallel to the transverse axis (y) of the aircraft (1) and extends at right angles to the longitudinal beams (7.1, 7.2), the crossbeam (8) connects to a plurality of lifting units (3); The longitudinal beams (7.1, 7.2) are spaced apart from each other along the transverse axis (y) of the aircraft (1), and are rotatably connected to the fuselage (2) via at least one corresponding first pivot joint designed to pivot the longitudinal beams (7.1, 7.2) about a corresponding first pivot axis to a pivot position (7.1', 7.2'), wherein the first pivot axis is parallel to the longitudinal axis (x); and The crossbeam (8) is rotatably connected to the fuselage (2) via at least one second pivot joint for pivoting the crossbeam (8) about a second pivot axis to a pivot position (8'), wherein the second pivot axis is parallel to the vertical axis (z) of the aircraft (1).

2. The aircraft according to claim 1, wherein, The lifting unit (3) includes at least one corresponding rotor (3b) arranged in a common rotor plane, which is used at least for the longitudinal beams (7.1, 7.2) and / or for the crossbeam (8).

3. The aircraft according to claim 1 or 2, wherein, The crossbeam (8) is arranged above the longitudinal beams (7.1, 7.2) along the vertical axis (z) of the aircraft (1).

4. The aircraft according to claim 1, wherein, The longitudinal beams (7.1, 7.2) have multiple branches, and the lifting unit (3) is located at the corresponding bifurcation point of the branch and / or the end point of the branch.

5. The aircraft according to claim 1, wherein, The crossbeam (8) has multiple branches, and the lifting unit (3) is located at the corresponding bifurcation point of the branch and / or the end point of the branch.

6. The aircraft according to claim 1, wherein, Any rotor (3b) included in the lifting unit (3) can be connected to a rotational position relative to the longitudinal beam (7.1, 7.2) and the crossbeam (8), respectively.

7. The aircraft according to claim 6, wherein, The rotor (3b) located on the longitudinal beams (7.1, 7.2) can be connected such that its blades are oriented parallel to the longitudinal beams (7.1, 7.2), while the rotor (3b) located on the crossbeam (8) can be connected such that its blades are oriented parallel to the crossbeam (8) and transverse to the longitudinal beams (7.1, 7.2).

8. The aircraft according to claim 1, wherein, The aircraft (1) includes a plurality of connecting elements (5) for interconnecting one of the plurality of longitudinal beams (7.1, 7.2) with the crossbeam (8) via each of the connecting elements (5), or interconnecting one of the plurality of longitudinal beams (7.1) with another of the plurality of longitudinal beams (7.1, 7.2).

9. The aircraft according to claim 8, wherein, The specified connecting element (5) is rotatably connected to one of the plurality of longitudinal beams (7.1, 7.2) or to the crossbeam (8) via a third pivot joint (5b), the third pivot joint (5b) being designed to rotate the connecting element (5) to abut against one of the plurality of longitudinal beams (7.1, 7.2) or the crossbeam (8).

10. The aircraft according to claim 9, wherein, The aircraft (1) includes a quick-release mechanism (5a) for fixing the connecting element (5) to the crossbeam (8).

11. The aircraft according to claim 1, wherein, The longitudinal beams (7.1, 7.2) surround the fuselage (2) and the landing gear (2'') and / or payload (2') optionally connected to the fuselage (2) at their pivot positions (7.1', 7.2').

12. The aircraft according to claim 1, wherein, The aircraft (1) includes a flexible mechanism designed to support the pivoting motion of the longitudinal beams (7.1, 7.2).

13. The aircraft according to claim 1, wherein, The aircraft (1) includes a detachable beam (8) clamp that provides the second pivot joint.

14. The aircraft according to claim 1, wherein, The aircraft (1) includes a fixing device for securing the crossbeam (8) to the longitudinal beam (7.1, 7.2) at the pivot position.

15. The aircraft according to claim 1, wherein, The aircraft (1) comprises an even number of lift units, N=2k, where k∈N, and each lift unit has at least one rotor (3b), wherein: When N=6, the crossbeam (8) is provided with two lifting units (3) at its opposite ends, and each of the two longitudinal beams (7.1, 7.2) is provided with two lifting units (3) at its opposite ends. When N=8, four lifting units (3) are provided on the crossbeam (8), and two lifting units (3) are provided at each of the two longitudinal beams (7.1, 7.2) at their opposite ends. When N=18, six lifting units (3) are provided on the crossbeam (8), and each of the two longitudinal beams (7.1, 7.2) is provided with six lifting units (3).

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