ELECTRICALLY POWERED AIRCRAFT

Description

[0001] The invention relates to an electrically powered aircraft with the features according to the preamble of claim 1.

[0002] Conventional aircraft can typically be divided into two categories. First, there are airplanes with a pair of wings for aerodynamic flight, usually equipped with a forward-acting propulsion system. Second, there are helicopters with a hovering propulsion system that primarily acts along the vertical axis, enabling hovering flight. The latter allows takeoff and landing even in very confined spaces. Conventional airplanes have a significantly higher efficiency in aerodynamic flight but require considerable space and infrastructure for takeoff and landing.

[0003] Currently, there are various efforts to combine the advantages of one concept with those of another. Numerous studies and prototypes of aircraft exist that feature both wings for aerodynamic flight with a forward-acting propulsion system and a hovering propulsion system acting vertically. Most aircraft of this type share the common characteristic that the hovering propulsion system is derived from the concept of the so-called multicopter. Such a hovering system comprises several propellers arranged in four quadrants of a horizontal plane of the aircraft. Mechanically complex cyclic pitch control, as found in classic helicopters, is omitted. Instead, control is achieved through individual speed adjustment of each propeller.The propellers are designed to rotate alternately clockwise and counterclockwise, so that in the neutral state, a torque balance exists for the sum of all drive torques. Different power and speed control of the individual propellers allows for steering around all spatial axes, thus eliminating the need for a tail rotor as found in a conventional helicopter.

[0004] Such a concept is known, for example, from DE 10 2013 109 392 A1. The aircraft shown there has a front and a rear pair of wings, with a lifting propeller at each wingtip, modeled on a quadcopter. This is intended to enable hovering for takeoff and landing. During cruise flight, the aircraft is supported by its two pairs of wings, with an independent propulsion system providing thrust.

[0005] However, several difficulties remain unresolved, which play a significant role in practical implementation. Controlling the aircraft in hover using the various propellers is fundamentally possible, but becomes increasingly difficult with the growing size and mass of the aircraft. While the familiar multicopter control system has reached a considerable level of quality in model-making, such a system cannot be easily scaled up to a larger, manned aircraft. The high inertia of a manned aircraft is difficult to overcome using copter control methods, resulting in sluggish control response.

[0006] Another problem that remains unsatisfactorily resolved is the so-called transition, i.e., the change from hovering to aerodynamic flight and vice versa. After vertical takeoff and with increasing forward speed, the lift engine must keep the aircraft airborne until the wings generate the necessary lift. Conversely, during the transition from cruise flight to hovering for landing, the aircraft, with decreasing airspeed, approaches its stall speed, at which the airflow over the wings separates. To prevent the aircraft from dropping too low, the lift engine must provide the corresponding lift immediately and without delay. Therefore, both during the transition from hovering to cruise flight and the reverse transition, there is a speed range in which the aerodynamic lift and the lift from the lift engine overlap.Unless special measures are taken, undefined flight conditions may occur which cannot be easily controlled by the pilot.

[0007] An aircraft known from US patent 2018 / 105267 A1 is based on a conventional aerodynamic concept with a main wing and a tail assembly located behind it. In addition to a forward-acting propulsion system, it has several lift propellers arranged in pairs in front of and behind the main wing, with their propellers having inclined axes of rotation. EP 3 098 161 A1 describes an aircraft in a canard configuration with a total of four laterally inclined lift propellers. US patent 2019 / 0127056 A1 also discloses an aircraft in a canard configuration, featuring a total of six lift propellers with laterally inclined axes of rotation. The two central lift propellers are also inclined forward.

[0008] The invention is based on the objective of further developing a generic aircraft in such a way that good flight performance and clean controllability are achieved over the entire speed range, including hovering flight.

[0009] This problem is solved by an aircraft having the features of claim 1.

[0010] The invention is based on the understanding that conventional copter control, particularly around the vertical axis, relies on the torque balance of the individual propellers or their associated lift motors. In steady hovering flight, the sum of all drive torques acting on the propellers is zero and therefore balanced. Consequently, the aircraft as a whole experiences no torque around the vertical axis, and thus remains in its existing longitudinal orientation. However, if the counter-rotating propellers are operated with different drive torques (different acceleration or deceleration, maintaining different rotational speeds of the counter-rotating propellers), the previously existing torque balance is disrupted. This results in a net torque around the vertical axis, which leads to a yaw movement of the aircraft.

[0011] However, the net yaw moment generated according to the above concept is not as pronounced from a certain size of the aircraft, and especially in a manned design, as is required for the control responsiveness desired by the pilot.

[0012] According to the invention, it is therefore provided that the net yaw moment generated from the drive torques is supplemented by an additional yaw moment component. For this purpose, the invention provides that, in a laterally opposed pair of propellers, the corresponding axes of rotation are inclined laterally in opposite directions with respect to the vertical axis and meet at an intersection point. In the case of a pair of propellers located forward of the center of gravity, this intersection point of the corresponding axes of rotation lies forward of the intersection point of the axes of rotation of a pair of propellers located aft of the center of gravity. Furthermore, it is provided that within a quadrant, all propellers located therein have the same direction of rotation about their respective axes of rotation.Advantageously, while maintaining the aforementioned concept for a fail-safe design, several and in particular two propellers, driven about their axis of rotation and with the same direction of rotation, are provided in each quadrant, wherein the points of intersection of the axes of rotation of the pairs of propellers located in front of the center of gravity are in front of the points of intersection of the axes of rotation of the pairs of propellers located behind the center of gravity.

[0013] The inclination of the axes of rotation according to the invention essentially retains the aforementioned concept of torque control, but adds an additional torque component as follows: The inclination of the axes of rotation initially causes the thrust of each propeller to act not only vertically upwards, but also proportionally outwards from the longitudinal axis. If a pair of laterally opposed propellers, for example positioned in front of the center of gravity, are operated at different speeds, a net lateral force is generated while the average lift remains constant. The same applies to a pair of laterally opposed propellers behind the center of gravity, although in this case the resulting net lateral force acts in the opposite lateral direction.Due to the positioning of the intersection points of the axes of rotation defined above, the opposing lateral forces form a counteracting force couple with a resulting torque about the vertical axis. This resulting torque adds to the torque already present, which results from the different drive torques of the propellers. Overall, this provides a significantly increased total torque about the vertical axis, which can be used for precise yaw control of even larger and therefore more sluggish aircraft, particularly in a manned version. Practical flight tests have shown that the inventive concept of axis tilting provides a yaw moment range that is sufficient for controlling a larger, more sluggish, and even manned aircraft, well beyond the scale of a model.

[0014] In a further aspect of the invention, it should be emphasized that all propellers in diagonally opposite quadrants have the same direction of rotation. All propellers within a quadrant support each other and, in the event of a propeller failure, can ensure load-bearing capacity and controllability as a "fail-safe" concept. Furthermore, the control of the aircraft around the various spatial axes (yaw, roll, and pitch control) is decoupled: All propellers in two quadrants located on one diagonal can be operated with increased torque and thrust when rotating in the same direction, while all other propellers in the quadrants on the other diagonal rotate in the opposite direction but with correspondingly reduced torque and thrust. This results in a pure yaw moment without any change in the total thrust.of the total lift, and also without a resulting roll or pitch moment. A roll or pitch moment, however, can be generated if all propellers on the same side of the longitudinal or transverse axis are operated with increased torque and thrust, while all propellers on the opposite side of the longitudinal or transverse axis are operated with reduced torque and thrust. The result is a roll or pitch moment, causing the aircraft to roll or pitch, while the total torque about the vertical axis remains zero, meaning no yaw motion occurs. The total thrust, and therefore the total lift, also remains unchanged.

[0015] It can be advantageous to tilt the helicopter propellers inwards. Preferably, however, all axes of rotation are tilted outwards, so that the axes of rotation in the vertical direction have a greater distance from the longitudinal center plane of the aircraft at the top than further down. In this configuration, the direction of rotation of all helicopter propellers is counterclockwise in the forward right quadrant and in the rear left quadrant, and clockwise in the forward left quadrant and in the rear right quadrant. This ensures, firstly, that the torque components generated at varying rotational speeds are complementary with the correct sign. Secondly, the lines of intersection of the rotor planes of rotation with the longitudinal center plane of the aircraft are raised sufficiently to lie above a pilot or passenger compartment located in the fuselage. This represents a safety aspect relevant for certification.

[0016] The inclination of the rotation axes relative to the vertical axis can be specified by a lateral angle, which is advantageously in the range of 3° to 8° and particularly 5°. This has proven suitable for achieving the aforementioned effects without a significant loss of vertical lifting force, a fact that has also been confirmed in flight tests.

[0017] Alternatively or additionally to the aforementioned lateral angle, a forward tilt of the propeller axes relative to the vertical axis by a pitch angle can also be advantageous, with the pitch angle being in the range of 2° to 5° and particularly around 3.5°. In hovering flight, when the propeller axes are naturally approximately vertical, this results in the aircraft being slightly tilted backward along its longitudinal axis, which is not perceived as particularly bothersome by the pilot or passengers. If, however, forward acceleration of the aircraft is to be initiated, a forward tilt of the propellers is necessary. For this purpose, the aircraft is tilted forward from its backward-tilted resting position. Due to the relative tilt angle of the propeller axes, the desired tilt of the axes is achieved even when the longitudinal axis of the aircraft is approximately horizontal.The aircraft accelerates forward in an approximately horizontal orientation. This facilitates the transition to aerodynamic cruise flight, as the wings are subjected to the airflow at approximately the desired angle of attack during forward flight. Conversely, the termination of aerodynamic flight due to the reduction in airspeed and increase in angle of attack causes the helicopter propellers to automatically return to their approximately vertical hovering orientation, without generating the desired forward thrust from the helicopter propellers that would impede deceleration.

[0018] The propellers move within their respective rotation circles. In an advantageous embodiment, the propellers are arranged such that rotation circles positioned one behind the other in the longitudinal direction have a vertical offset relative to each other in the direction of the axes of rotation, and these rotation circles also partially overlap. This allows the use of propellers with a comparatively large diameter at a comparatively low rotational speed. This results in a comparatively low power requirement for hovering flight, while maintaining high efficiency. Economic efficiency and, in particular, suitability for electric motor propulsion are improved.

[0019] The propellers each have two blades, which are advantageously arranged at a conical angle with respect to their axes of rotation, and where the conical angle is less than 180°. Centrifugal and lift forces acting on the propeller blades thus cancel each other out to such an extent that a substantially radial load occurs, avoiding bending moments. Accordingly, the propellers can be made thin and lightweight. Despite this, deformation under load is minimal, allowing for close spacing without risk of collision. The conical angle also increases the clearance between the propellers and the longitudinal spars and wingtips.

[0020] According to the invention, the aircraft is designed as a tandem-wing aircraft with a front pair of wings and a rear pair of wings, wherein the front pair of wings and the rear pair of wings have substantially the same wingspan. The wingtips of the front wings are connected to corresponding wingtips of the rear wings by means of a longitudinal strut, wherein the lifting propellers, in particular including their lifting motors, are mounted on the longitudinal struts.

[0021] The aforementioned configuration allows the lift propellers to be positioned longitudinally independently of the actual position of the wingtips, so that a common center of lift for all lift propellers can be brought at least close to the aircraft's design center of gravity. This results in all lift propellers being subjected to approximately the same load during hovering. All lift motors retain their control reserves in steady hovering without having to constantly compensate for an off-center center of gravity. The design center of gravity itself is located sufficiently far forward of the aerodynamic neutral point, enabling stable, straight-line flight.The same trim conditions apply, at least approximately, to both straight and aerodynamic flight and hovering, allowing for transitions between these two flight states without significant trim changes. Furthermore, the longitudinal spars, together with the wings, form a frame that stiffens the load-bearing structure, particularly with regard to torsion.

[0022] In a further advantageous development of the aforementioned tandem-wing concept, the forward pair of wings has a first mean chord, while the rear pair of wings has a second mean chord, with the first mean chord being in the range of 30% to 40% of the second mean chord. This distribution of mean chords allows the aerodynamic neutral point and the center of gravity, positioned at a sufficient safety margin in front of it, to be located approximately at the geometric center of the aircraft. The center of gravity and the overall center of lift of all propellers can thus be effectively aligned.

[0023] In a preferred embodiment, the leading pair of wings is positioned higher along the vertical axis than the trailing pair. This ensures that, during slow flight at high angles of attack and even in a stall, the trailing pair never encounters any wake turbulence that may form from the leading pair of wings. This allows for recovery using only aerodynamic means and without the use of the elevator.

[0024] Various conventional means, such as rudders, elevators, and ailerons, are suitable for aerodynamic control. Preferably, the aircraft is equipped with elevons for aerodynamic pitch and roll control, which are arranged particularly on the trailing pair of wings. Here, the elevator and ailerons are combined into a single control surface. A division of the control surfaces along the wingspan is not necessary. This ensures that even with a shorter wingspan, sufficiently large control surfaces with corresponding control effects are available. Overall, this results in a clean aerodynamic design with few undesirable gaps and simple kinematic control. The elevons are moved or controlled in the conventional manner using a control stick, such that pulling or pulling the stick results in a controlled movement.Pushing the control stick results in a corresponding elevator deflection, and a lateral movement of the control stick results in an opposing aileron deflection of the elevons. For this purpose, two pushrods, preferably arranged in a V-shape directly adjacent to the control stick, are connected to each elevon via bell cranks and other transmission elements, particularly Bowden cables. Without any additional kinematic complexity, the V-shape acts as a mechanical or kinematic mixer, converting a pull or push of the control stick into a corresponding elevator deflection and a left-right movement of the control stick into an opposing aileron deflection of the elevons. By adjusting the bell crank angles, differential elevon deflections can also be achieved without additional complexity to reduce or even eliminate yawing and roll moment.

[0025] In a preferred embodiment, the elevons are equipped with spades. Due to their size, the elevons can generate high control forces, especially at high speeds, which may need to be reduced for purely manual control. As compensating surfaces arranged in front of the elevons' axis of rotation, the elevons reduce the required manual forces for generating aileron and elevator deflection. Furthermore, it has been shown that, in aerodynamic flight, the elevons tend to drift upwards from their design-defined neutral position in the direction of a common, synchronized elevator deflection due to the aerodynamic forces acting upon them. To avoid the need for constant counter-steering in this regard, the spades are advantageously trimmed such that the elevons assume a substantially neutral trim position with respect to pitch control during aerodynamic flight.Elevons adjusted in this way result in a manual elevator trim during steady flight, eliminating the need for constant counter-steering. In another respect, the spades counteract flutter. Finally, especially in conjunction with the spades, counterweights can be attached and positioned to achieve the most complete mass balance of the elevons possible. This also counteracts flutter. Furthermore, the counterweights generate neutral control moments even during takeoff with little or no airflow, allowing the elevons to be brought into and held in their neutral position without manual effort.

[0026] An embodiment of the invention is described in more detail below with reference to the drawing. The drawing shows: Fig. 1 shows a perspective view of an aircraft designed according to the invention with front and rear wings for aerodynamic flight and with lifting propellers to form a hovering propulsion system; Fig. 2 shows a top view of the aircraft according to the invention. Fig. 1 with details on the positioning of the front and rear wings and the lifting propellers relative to a design center of gravity, Fig. 3 in a front view of the aircraft according to the Figs. 1 and 2 with details on the vertical spacing of the wings and the lateral tilt of the helicopter propellers, Fig. 4 in a side view of the aircraft according to the Figs. 1 to 3with details of a tilt angle of the lifting propellers and the relative positioning of the intersection points of the axes of rotation, Fig. 5 in a cross-sectional view of one of the rear wings with an elevon for aerodynamic roll and pitch control as well as with a spade mounted on the elevon, and Fig. 6 in a perspective detail view details of a control mechanism for controlling the elevons according to the Fig. 1, 2 and 5 .

[0027] Fig. 1Figure 1 shows a perspective view of a manned aircraft designed according to the invention, in this case a single-seat aircraft. The aircraft has a fuselage 1 with a canopy located approximately in the center, which covers the cockpit of a single pilot while providing the pilot with an unobstructed view. However, a two- or multi-seat version, or even an unmanned version, is also possible. The fuselage 1 is designed as an aramid sandwich safety cell. The remaining structural components of the aircraft consist primarily of carbon fiber laminate.

[0028] Extending laterally from the fuselage 1 are a front pair of wings 2 and a rear pair of wings 3. During aerodynamic flight of the aircraft in a forward direction u, which essentially corresponds to a longitudinal axis x of the aircraft, the front wings 2 and the rear wings 3 provide the necessary lift due to the airflow around them. Propulsion in the forward direction u is provided by an electric flight propulsion system 10 (not shown in detail) in the form of a motor with a propeller, of which only a suggested propeller circle is shown here for clarity. The aircraft also features an electric hovering propulsion system for hovering flight, which acts primarily in the direction of a vertical axis z of the aircraft. The hovering propulsion system includes various lifting propellers 11, 12, 13, 14, which can be driven to rotate about their respective axes of rotation.In the illustrated embodiment, each lifting propeller 11, 12, 13, 14 has its own lifting motor 25. However, a central drive or motor is also possible. For the sake of simplicity, the lifting motors 25 are only recognizable here by their aerodynamic fairings. The lifting motors 25 and the flight propulsion motor 10 are electric motors equipped with magnets in a Halbach array configuration for superior efficiency. They are powered by specially developed, adapted sine-pulse power controllers from electric batteries (not shown) carried in the aircraft. The resulting aircraft is capable of hovering using its hovering propulsion system and aerodynamic flight using its wings 2, 3 and the flight propulsion system 10.

[0029] At the rear end of the fuselage 1, relative to the forward direction u, a conventional vertical stabilizer 36 with a rudder 37 is mounted, providing stabilization and yaw control around the vertical axis z during aerodynamic forward flight. Along their trailing edges, the rear wings 3 are each equipped with an elevon 7, which serves as an aerodynamic control surface. Alternatively or additionally, the control surfaces, or elevons 7, can also be located on the forward wings 2. The elevons 7 extend continuously across the entire half-span of each respective rear wing 3. The term "elevon" is a combination of "elevator" for horizontal stabilizer and "aileron" for aileron, and refers to aerodynamic control surfaces for combined pitch and roll control.

[0030] For takeoff, landing, and ground taxiing, the aircraft is equipped with a wheeled landing gear, designed here as a classic bipod with an additional rear tailwheel 38. However, other landing gear configurations may also be suitable. In any case, the landing gear allows takeoff and landing in purely aerodynamic flight mode with sufficient forward speed, utilizing the lift generated by the front and rear wings 2, 3. If the option of aerodynamic takeoff or landing is not desired, the landing gear can also be replaced by a suitable frame or by skids modeled on a helicopter.

[0031] Fig. 2 The aircraft is shown in a top view. Fig. 1with further details. The aircraft extends along the aforementioned longitudinal axis x, which also essentially determines the forward direction u, i.e., apart from variations in the angle of attack and yaw angle. A transverse axis y runs perpendicular to the longitudinal axis x, defining the wingspan direction of the front and rear wings 2 and 3. The vertical axis z runs perpendicular to both the longitudinal axis x and the transverse axis y, passing through the intersection point. The aircraft has a design center of gravity CG, where the origin of the coordinate systems of the longitudinal axis x, the transverse axis y, and the vertical axis z is located at the design center of gravity CG. From the design center of gravity CG, the aircraft can have an actual center of gravity that lies a limited distance forward or aft of the design center of gravity CG within a permissible range.The pilot sits almost exactly at the design center of gravity (CG), so different pilots of varying weights have no influence on the actual center of gravity. During active operation, variations in the actual center of gravity are primarily determined by any additional load carried (payload, luggage). Of course, within the scope of the invention, it is also possible that the pilot and any passengers are not positioned directly at the design center of gravity (CG), so changes in the load configuration can also lead to variations in the actual center of gravity. In such a case, it must be ensured that the actual center of gravity lies within a permissible range.

[0032] The front wings 2 and the rear wings 3 have the same wingspan in the direction of the lateral axis y, but differ in their chord. The front pair of wings 2 has a first mean chord l1, while the rear pair of wings has a second mean chord l2. In the preferred embodiment shown, the first mean chord l1 is only about 30% to 40% of the second mean chord l2. This, in conjunction with the other aerodynamic parameters of the front and rear wings 2, 3, such as the angle of incidence difference, the moment coefficients of the selected airfoils, and the like, results in a position of the aerodynamic neutral point NP sufficiently far behind the center of gravity CG, so that sufficient stability and controllability about the lateral axis y are always available during aerodynamic flight.

[0033] In addition to the aforementioned details regarding the aerodynamic and flight-mechanical design, details for the hovering propulsion system are also important. These details are shown in the top view according to... Fig. 2It is evident that the longitudinal axis x and the transverse axis y define four quadrants I, II, III, IV of the aircraft in the plane they span. In each of these quadrants, the hovering drive has at least one propeller 11, 12, 13, 14, each driven to rotate about a corresponding axis of rotation 15, 16, 17, 18. In the preferred embodiment shown, each quadrant I, II, III, IV contains several, and here exactly two, propellers 11, 12, 13, 14, each driven to rotate in the same direction about their respective axes of rotation 15, 16, 17, 18 in a fail-safe configuration. With respect to the longitudinal direction 8, the respective propellers 11, 12, 13, 14 are positioned laterally opposite each other in pairs. In other words, each individual propeller 11 of the first quadrant I is opposite a propeller 12 of the second quadrant II, while each propeller 13 of the third quadrant III is opposite a propeller 14 of the fourth quadrant IV.All propellers 11 located within the first quadrant (I) located in the forward direction u, the front right quadrant, have the same direction of rotation, namely counterclockwise when viewed from above, as indicated by arrow 26. The same applies to all propellers 13 located in the third quadrant (III) located in the forward direction u, the rear left quadrant. All propellers 12 located in the second quadrant (II) located in the forward direction u, the front left quadrant, rotate in the opposite direction, i.e., clockwise when viewed from above, as indicated by arrow 27. The same applies to all propellers 14 located in the fourth quadrant (IV) located in the forward direction u, the rear right quadrant.In other words, all the propellers 11, 13 located in diagonally opposite quadrants I, III have the same counterclockwise direction of rotation, while all the propellers 12, 14 located in diagonally opposite quadrants II, IV have the same clockwise direction of rotation.

[0034] The lifting propellers 11, 12, 13, 14 each have two propeller blades 44, 45, although a different number of propeller blades may also be advantageous. They are so-called "fixed-pitch" propellers with a fixed pitch. It follows that their thrust is adjusted solely by the rotational speed. The lifting propellers 11, 12, 13, 14 differ only in their adaptation to the respective direction of rotation and are otherwise identical, particularly with regard to diameter and pitch. The lifting motors 25 Fig. 1and thus the available drive power of all propellers 11, 12, 13, 14 is the same.

[0035] The propellers 11, 12, 13, 14 are positioned in the top view shown as mirror images of each other with respect to both the longitudinal axis x and the transverse axis y, such that their geometric center lies at the intersection of the longitudinal axis x and the transverse axis y, and thus at the center of gravity CG. With the drive power of all propellers 11, 12, 13, 14 adjusted to the same level, their common center of thrust coincides with the center of gravity CG. The total thrust of all propellers 11, 12, 13, 14 keeps the aircraft hovering, with a moment equilibrium about the longitudinal axis x, the transverse axis y, and the vertical axis z in the described initial state. Consequently, the aircraft does not exhibit any pitching, rolling, or yaw movements.

[0036] Starting from this, a roll motion about the longitudinal axis x is controlled or generated by operating the propellers 11, 14 on one side of the longitudinal axis x at a speed that differs from the speed of the propellers 12, 13 on the opposite side of the longitudinal axis x. This speed difference can be adjusted such that the total thrust and thus the lifting force remains constant, while a roll moment about the longitudinal axis x is generated. Due to the opposing directions of rotation of the propellers 11, 13 in the first and diagonally opposite third quadrants I, III on the one hand, and the propellers 12, 14 in the second and diagonally opposite fourth quadrants II, IV on the other, no net total moment about the vertical axis z is generated despite the changes in speed, so that there is practically no coupling between the roll control and the yaw control. The same principle applies analogously to a control or...the generation of a pitching motion about the transverse axis y by operating the helicopter propellers 11, 12 on one side of the transverse axis y at a rotational speed that differs from the rotational speed of the helicopter propellers 13, 14 on the opposite side of the transverse axis y.

[0037] A yaw movement about the vertical axis z is controlled or generated proportionally by operating the propellers 11, 13 in the first and diagonally opposite third quadrants I, III, and the propellers 12, 14 in the second and diagonally opposite fourth quadrants II, IV, with different drive torques. The different drive torques of the counterclockwise rotating propellers 11, 13 and the clockwise rotating propellers 12, 14 generate a net total torque about the vertical axis z in the form of a yaw torque, which leads to a yaw movement. According to the invention, this net total torque, which can be accessed when using the hover drive, is further supplemented by an additional torque component resulting from a Fig. 2The total yaw moment about the vertical axis z can be adjusted in this way without changing the total lift or causing unwanted roll or pitch moments. Further details are given below in connection with... Fig. 3 described.

[0038] From the overall view of Figs. 1 and 2It becomes clear that the lifting propellers 11, 12, 13, 14 with their axes of rotation 15, 16, 17, 18 and with their lifting motors 25 are not located directly at the outer wingtips 4, 5 of the wings 2, 3. Rather, the wingtips 4 of the forward wings 2 are connected to the wingtips 5 of the rear wings 3 on the same side by means of a longitudinal beam 6 running approximately parallel to the longitudinal axis x. The longitudinal beams 6 serve to accommodate the lifting motors 25 and to support the lifting propellers 11, 12, 13, 14 for a rotating drive about their respective axes of rotation 15, 16, 17, 18, and allow them to be positioned independently of the actual position of the wingtips 4, 5. Furthermore, together with the wings 2, 3, they form a substantially rectangular frame, which contributes to the torsional stiffness of the entire structure.

[0039] Fig. 3shows a front view of the aircraft after the Figs. 1 and 2 . Here as well as in the side view after Fig. 4 It is immediately apparent that, with the longitudinal axis x horizontal, the front pair of wings 2 is higher in the direction of the vertical axis z than the rear pair of wings 3. This prevents any wake vortex potentially generated by the front wings 2 from reaching the rear wings 3.

[0040] Furthermore, it follows from the Fig. 3The laterally opposed pair of front propellers 11, 12, with their associated axes of rotation 15, 16, are inclined laterally by an angle δ in opposite directions with respect to the vertical axis z. Both axes of rotation 15, 16 are inclined upwards and outwards by the same amount of angle δ, which is preferably in the range of 3° to 8° and is approximately 5° in the illustrated embodiment. This lateral inclination causes the axes of rotation 15, 16 to meet or cross centrally below the aircraft. The same applies to all other pairs of propellers 11, 12 and 13, 14. Fig. 2 .

[0041] Fig. 4 shows a side view of the aircraft after the Figs. 1 to 3In cruise flight, the longitudinal axis x is approximately horizontal and essentially coincides with the forward direction u. At slower airspeeds and the associated higher angle of attack, the nose of the aircraft may point slightly upwards, so that the longitudinal axis x is angled a few degrees above the forward direction u. Conversely, in high-speed flight, a slight negative angle of attack of the longitudinal axis x relative to the forward direction u may occur. In any case, it is evident that the axes of rotation 15, 16, 17, 18 of the helicopter propellers 11, 12, 13, 14 ( Fig. 2 ) are inclined forward about the vertical axis z by a tilt angle ε. The pitch angle ε is advantageously in the range of 2° to 5° and is approximately 3.5° in the illustrated embodiment. Furthermore, the following results from the combination of Figures 2 and 4that the propellers 11, 12, 13, 14, more precisely, that the tips of their propeller blades 44, 45, move in corresponding circles of rotation 21, 22, 23, 24. Circles of rotation 21, 24 and circles of rotation 22, 23, which are consecutive along the longitudinal axis x, have a vertical offset relative to each other in the direction of their axes of rotation 15, 16, 17, 18, and overlap proportionally. Despite the overlap, the vertical offset prevents a collision of the propellers 11, 12, 13, 14. Furthermore, the propeller blades 44, 45 are arranged at a conical angle φ with respect to the axes of rotation 15, 16, 17, 18, where the conical angle φ is less than 180°. In other words, the propeller blades 44, 45 are angled slightly upwards. The cone angle φ reduces the loads and deformations acting on the propeller blades 44, 45 and ensures sufficient clearance from the longitudinal beams 6 and the wingtips 4, 5 of the wings 2, 3.

[0042] The overview of Figs. 3 and 4 It can be seen that the axes of rotation 15, 16 of the front propellers 11, 12 intersect at front intersection points 19, while the axes of rotation 17, 18 of the rear propellers 13, 14 intersect analogously at rear intersection points 20. (See side view.) Fig. 4 It can be deduced that the forward intersection points 19 of the associated forward axes of rotation 15, 16 lie in front of the rear intersection points 20 of the rear axes of rotation 17, 18. This allows the yaw control of the aircraft about the vertical axis z described above to be supplemented as follows: First, in Fig. 3It can be seen that the two laterally opposed propellers 11, 12 generate the same force F, directed obliquely outwards and upwards, at the same rotational speed. Due to the inclination of the axes of rotation 15, 16 by the lateral angle δ, these forces F can be resolved into vertical force components Fz and horizontal force components Fy. Since both horizontal force components Fy are equal in magnitude and opposite in direction, they cancel each other out to zero in the initial state.

[0043] Starting from the aforementioned initial state, for yaw control the propellers 11, 13 of the first and third quadrants I, III are operated at a speed that differs from the speed of the propellers 12, 14 of the second and fourth quadrants II, IV. The lateral tilt of the forward axes of rotation 15, 16 by the lateral angle δ ( Fig. 3This leads to the situation that while the sum of the vertical force components Fz and thus the total lift remains the same, the two horizontal force components Fy are no longer equal in magnitude and therefore no longer cancel each other out. Rather, the front propellers 11, 12, which operate at different speeds, generate a resultant front lateral force Fyv. The rear propellers, here in Fig. 3 The propellers 13, 14, not shown but also operated at different speeds, generate a resultant rear lateral force F yh in an analogous manner, which together with the front lateral force F yv in Fig. 2is entered and acts in the opposite direction to the front lateral force F yv. The previously mentioned distance between the front intersection points 19 and the rear intersection points 20 converts these two opposing horizontal lateral forces F yh, F yv into a yaw moment about the vertical axis z. This yaw moment resulting from the horizontal lateral forces F yh, F yv adds to the sum of all drive torques acting about the axes of rotation 15, 16, 17, 18 described above, which no longer cancel each other out due to the different loads. The resulting total torque can be used for yaw control about the vertical axis z when using the floating drive.

[0044] From above Fig. 2It is still recognizable that the elevons 7 are each equipped with a so-called spade 9. A "spade" is an aerodynamic compensation surface that is attached to an aerodynamic control flap by means of a cantilever arm in such a way that it lies in front of the control flap's pivot axis and thereby reduces the flap moments generated during a control deflection and consequently the hand forces required for operation. Fig. 5Figure 1 shows a cross-sectional view of one of the rear wings 3 in the area of ​​such a spade 9. The aerodynamic profile of the rear wing 3, extending along a chord line 34, is visible in the cross-section. The wing 3 is fixed in the leading edge of the profile. In the trailing edge, it is designed as an aerodynamic control flap in the form of an elevon 7, the elevon 7 being pivotally attached to the fixed part by hinges and pivotally mounted about a pivot axis 30 perpendicular to the plane of the drawing.

[0045] A lever arm 33 is attached to the elevon 7, extending upwards from the elevon 7 and forwards in the forward direction u to just in front of the pivot axis 30. A spade 9 is attached to this outer, free end of the lever arm 33, and a counterweight 32 is attached in its immediate vicinity. Positioned in front of the pivot axis 30 in the forward direction u, the counterweight 32 at least partially compensates for the mass of the elevon 7 located behind the pivot axis 30, in this case even 100%. This mass balancing thus serves as a measure against rudder and wing flutter and contributes to more pleasant controllability. Finally, the mass balancing prevents the elevons 7 from tilting downwards under their own weight, allowing them to be moved effortlessly into their desired neutral position and held there even without aerodynamic flow.

[0046] The spade 9 can be designed as a flat plate. In the illustrated embodiment, it has a load-bearing aerodynamic profile. In any case, as already mentioned, the spade 9 contributes to reducing the flap moment acting on the elevon 7 about the pivot axis 30 and thus to reducing the manual forces required for control, as is known in the prior art when used on conventional ailerons. However, in the present application on the elevons 7, a further application is added. Unlike pure ailerons, the elevons 7 not only deflect in opposite directions, but also in the same direction when functioning as elevators. It has been shown that the aerodynamic air pressure distribution acting on the airfoil of the wings 3 tends to pull the elevons 7 upwards about their pivot axis 30 in the same direction as arrow 31. This has the effect of an unwanted elevator trim, which must be compensated for.In addition to reducing hand force, the spades 9 according to the invention are used to compensate for this unwanted elevator trim. For this purpose, they are set at a specific angle y relative to the airfoil 34 when the control surface is in neutral position. In aerodynamic flight, this, along with the optional spade profile on the spade 9, generates a force that acts as a moment about the pivot axis 30 via the lever arm 33. The aerodynamic moment that would otherwise occur at the elevon 7 in the direction of arrow 31 can thus be at least approximately completely compensated.

[0047] The angle of incidence y and the profile of the spades 9 are trimmed or selected such that the elevons 7 assume an essentially neutral trim position with respect to pitch control during aerodynamic flight and maintain this position in Fig. 5The neutral position shown should be maintained even without applying any hand force. The neutral position is generally, at least approximately, the position without flap deflection, i.e., with a neutral flap angle of approximately 0°, as shown in Fig. 5 As shown. If necessary, a different neutral position of the elevons 7 can also be selected by adjusting the pitch angles y to the spades 9. In a further aspect, the spades 9 also act as a measure against flutter.

[0048] Fig. 6 shows, in a perspective detail view, the essential elements of the hull 1 ( Fig. 1 ) arranged control mechanism for controlling the elevons 7 ( Fig. 1, 2 , 5 ). When controlling the aircraft in aerodynamic flight, control of the rudder 37 ( Fig. 1The controls are provided in the usual manner via pedals (not shown) for foot control. Roll and pitch control are also carried out in the usual manner by means of a conventional control stick 8. However, the use of elevons 7 requires a kinematic mixing of the movement components of the control stick 8 as shown in the illustration. Fig. 6 .

[0049] The control stick 8, which is normally operated by the pilot's hand, is gimbal-mounted in bearings 39 and 40, as is customary, so that it can pivot about the longitudinal axis x and the lateral axis y. Pivoting the control stick 8 about the longitudinal axis x is intended to result in a roll movement of the aircraft, for which the elevons 7, in their function as ailerons, must deflect in the opposite direction. Pivoting the control stick 8 about the lateral axis y (pushing or pulling) is intended to result in a pitch movement of the aircraft, for which a corresponding deflection of the elevons 7, in their function as elevators, is required. Two pushrods 41 are provided for the kinematic mixing of these movements. These pushrods are located directly adjacent to the control stick 8 and are pivotally attached to it below the bearings 39 and 40.Starting from the lower end of the control stick 8, the pushrods 41 extend rearward in a V-shape, enclosing an opening angle of approximately 90° between them. At their rear, outer ends, the pushrods 41 are pivotally connected to bell cranks 42, which in turn are pivotally mounted on a common base plate 38. Bowden cables 43 extend from these bell cranks 42, each acting on one of the elevons 7.

[0050] The V-shaped arrangement of the pushrods 41 means that when the control stick 8 is moved about the lateral axis y, both pushrods 41 are moved in the same direction. This movement is then transmitted via the bellcranks 42 and the Bowden cables 43, resulting in a corresponding deflection of both elevons 7 in the direction of either a "pull" or a "push" action. The elevons 7 thus function as elevators for aerodynamic pitch control. Conversely, when the control stick 8 is moved about the longitudinal axis x, the pushrods 41 are moved in the opposite direction. This movement is transmitted via the bellcranks 42 and the Bowden cables 43, resulting in a corresponding deflection of both elevons 7, thus functioning as ailerons for aerodynamic roll control. The differential aileron deflection can also be adjusted by changing the angle of the bellcranks 42.

Claims

1. Aircraft which extends along a longitudinal axis (x), a transverse axis (y) and a vertical axis (z), wherein the longitudinal axis (x), the transverse axis (y) and the vertical axis (z) intersect in a centre of gravity of construction (CG), comprising at least one pair of wings (2, 3) for an aerodynamic flight in a forward direction (u) predefined by the longitudinal axis (x), a flight drive (10) acting mainly in the forward direction (u) for the aerodynamic flight, and a hover drive acting mainly in the direction of the vertical axis (z) for a hover flight, wherein the longitudinal axis (x) and the transverse axis (y) delimit in the plane defined by them four quadrants (I, II, III, IV) of the aircraft from one another, and wherein the hover drive in each quadrant (I, II, III, IV) has at least one propeller (11, 12, 13, 14) each which is able to be rotationally driven about its axis of rotation (15, 16, 17, 18), wherein the propellers (11, 12, 13, 14) are laterally opposite one another in pairs in terms of the longitudinal direction (x), wherein in the case of a laterally opposite pair of propellers (11, 12; 13, 14) the associated axes of rotation (15, 16; 17, 18) in terms of the vertical axis (z) are inclined in opposite directions towards the side and meet at an intersection point (19, 20), wherein in the case of a pair of propellers (11, 12) lying in front of the centre of gravity of construction (CG), the intersection point (19) of the associated axes of rotation (15, 16) lies in front of the intersection point (20) of the axes of rotation (17, 18) of a pair of propellers (13, 14) lying behind the centre of gravity of construction (CG), and wherein within a quadrant (I, II, III, IV) all propellers (11, 12, 13, 14) located therein have the same direction of rotation around their respective axis of rotation (15, 16, 17, 18), characterized in that the aircraft is designed as a tandem plane having a front pair of wings (2) and having a rear pair of wings (3), wherein the front pair of wings (2) and the rear pair of wings (3) have a substantially identical span, wherein wing ends (4) of the front wings (2) are in each case connected together to assigned wing ends (5) of the rear wings (3) by means of a longitudinal support (6), and wherein the propellers (11, 12, 13, 14), in particular including their lift motors (25), are mounted on the longitudinal supports (6).

2. Aircraft according to claim 1, characterized in that provided in each quadrant (I, II, III, IV) are in each case a plurality of, and in particular in each case two, propellers (11, 12, 13, 14) which are able to be rotationally driven in the same direction about their axis of rotation (15, 16, 17, 18), wherein the intersection points (19) of the axes of rotation (15, 16) of the pairs of propellers (11, 12) lying in front of the centre of gravity of construction (CG) lie in front of the intersection points (20) of the axes of rotation (17, 18) of the pairs of propellers (13, 14) lying behind the centre of gravity of construction (CG).

3. Aircraft according to claim 1 or 2, characterized in that, in terms of the forward direction (u), in the front right quadrant (I) and in the rear left quadrant (III) the direction of rotation of all propellers (11, 13) is anti-clockwise, in that, in terms of the forward direction (u), in the front left quadrant (II) and in the rear right quadrant (IV) the direction of rotation of all propellers (12, 14) is clockwise, and in that all axes of rotation (15, 16, 17, 18) are inclined outwards.

4. Aircraft according to one of claims 1 to 3, characterized in that the axes of rotation (15, 16, 17, 18) in terms of the vertical axis are inclined towards the side by a lateral angle (δ), wherein the lateral angle (δ) is in a range from 3° to 8°, and is in particular approximately 5°.

5. Aircraft according to one of claims 1 to 4, characterized in that the axes of rotation (15, 16, 17, 18) of the propellers (11, 12, 13, 14) in terms of the vertical axis (z) are tilted forward by a tilt angle (ε), wherein the pitch angle (ε) is in a range from 2° to 5°, and is in particular approximately 3.5°.

6. Aircraft according to one of claims 1 to 5, characterized in that the propellers (11, 12, 13, 14) move in associated rotation circles (21, 22, 23, 24), wherein rotation circles (21, 24; 22, 23) that are successive in the direction of the longitudinal axis (x) have a vertical offset in the direction of the axes of rotation (15, 16, 17, 18) and have a proportional overlap.

7. Aircraft according to one of claims 1 to 6, characterized in that the propellers (11, 12, 13, 14) have in each case in particular two propeller blades (44, 45), wherein the propeller blades (44, 45) in terms of the axes of rotation (15, 16, 17, 18) are disposed at a cone angle (φ), and wherein the cone angle (φ) is less than 180°.

8. Aircraft according to one of claims 1 to 7, characterized in that the front pair of wings (2) has a first mean wing depth (l1), in that the rear pair of wings (3) has a second mean wing depth (l2), and in that the first mean wing depth (l1) is in the range of 30% to 40% of the second mean wing depth (l2).

9. Aircraft according to one of claims 1 to 8, characterized in that the front pair of wings (2) in the direction of the vertical axis (z) is higher than the rear pair of wings (3).

10. Aircraft according to one of claims 1 to 9, characterized in that the aircraft for an aerodynamic pitch and roll control is provided with elevons (7).

11. Aircraft according to claim 10, characterized in that the elevons (7) are disposed on the rear pair of wings (3).

12. Aircraft according to claim 10 or 11, characterized in that the elevons (7) are moved by means of a joystick (8), wherein two push rods (41) directly adjacent to the joystick (8) are arranged in a V-shape and are in each case operatively connected to one elevon (7) each via deflection levers (42) and further transmission elements, in particular in the form of Bowden cables (43).

13. Aircraft according to one of claims 10 to 12, characterized in that the elevons (7) are provided with spades (9) which are trimmed in such a manner that the elevons (7) in the aerodynamic flight assume a substantially neutral trim position in terms of pitch control.

14. Aircraft according to one of claims 10 to 13, characterized in that a compensation mass (32) is disposed in the region of a spade (9).