eVTOL aircraft using large, variable-speed tilt rotors

Abstract

A VTOL aircraft, including: at least one wing mechanically coupled to a fuselage; at least four variable-speed tilt rotors, together designed and dimensioned to support a payload of at least 500 pounds; at least one first and at least one second motor, wherein the first and / or second motor is configured to drive a first of the at least four variable-speed tilting rotors; and at least one third and at least one fourth motor, wherein the third and / or fourth motor is configured to drive a second of the at least four variable-speed tilting rotors, wherein in a rotor-supported flight the at least four variable-speed tilt rotors are arranged together to lift the VTOL aircraft, wherein three of the at least four variable-speed tilt rotors are capable of providing at least 70% of the lift required to lift the VTOL aircraft carrying a payload of at least 500 pounds.

Description

priority

[0001] This application claims priority over preliminary US application No. 62 / 509666 (filed on May 22, 2017), preliminary US application No. 62 / 509674 (filed on May 22, 2017) and preliminary US application No. 62 / 656971 (filed on April 12, 2018), the disclosures of which are incorporated herein by reference. Field of invention

[0002] The invention is applied to aircraft that take off and land vertically. background

[0003] The background information contains details that may be useful for understanding the present invention. It does not constitute an admission that any of the information provided herein is prior art or relevant to the invention claimed herein, or that any publication expressly or implicitly referenced is prior art.

[0004] There is significant demand for electrically powered transportation. Due to the current limitations of rechargeable batteries (energy density - 220 Wh / kg - 300 Wh / kg, depth of discharge, charge / discharge rates, and lifespan issues), the general market entry sequence for battery-powered vehicles (apart from slow, short-range vehicles like golf carts) is as follows: • Automobiles – these vehicles are the easiest to convert to electric drive. Automobiles can accommodate heavy batteries, the battery discharge rate is relatively low, and operation can be safely stopped when the battery is depleted. • Motor glider - Here, an engine is used to launch an otherwise safe glider. • Fixed-wing aircraft for training - useful for short flights conducted from established airports with professional instructors, maintenance and management. • Private fixed-wing aircraft - the next easiest option thanks to roll take-off and landing with a high wing-based lift-to-drag ratio. • Electric VTOL (eVTOL) - more demanding due to the high power requirements for hovering, especially when efficient high-speed flight is also required.

[0005] eVTOL has become an even greater challenge as the market has shifted from “specialized transport aircraft” flying shorter distances (25-60 miles) between well-equipped terminals to “urban mobility aircraft” flying longer distances with at least one poorly equipped landing site. The need is illustrated by information published by Uber™ and is described here as the state of the art in the Fig. 1A and Fig. 1B is reproduced. Fig. 1A is a conceptual image of an emerging urban transportation market for the hybrid-electric, vertical take-off and landing (eVTOL) aircraft proposed by Uber®. Fig. 1B is a projected timeline for the development and operation of such an aircraft.

[0006] Safety and efficiency are perhaps the two most important factors in the development of eVTOL aircraft to serve this market. To achieve a high level of safety, much of the current state of the art focuses on aircraft that use six, eight, or even more independently operated rotors. If a single rotor of such an aircraft fails, the other rotors are likely to be able to perform a safe landing. Quadrotor aircraft are also not considered particularly fault-tolerant, as the failure of a single rotor can cause the aircraft to crash.

[0007] Several aircraft proposals and prototypes are currently being designed using this multi-rotor strategy. For example, the Fig. 2A and Fig. 2B Artistic renderings of a state-of-the-art Volocopter™ with 16 rotors, Fig. Figure 3 is a photo of a state-of-the-art 8-rotor Ehang™, and Fig. Figure 4 is a photograph of a state-of-the-art CityAirbus™ with eight rotors. However, all these designs are problematic because the rotors cannot tilt from the vertical lift position to the forward propulsion position, and there are no wings. This combination is extremely inefficient in forward flight, limiting the aircraft to relatively short ranges.

[0008] Currently, several eVTOL aircraft are being developed that continue to use the multi-rotor strategy but add a wing to improve forward flight efficiency. For example, Fig. Figure 5 is an artist's impression of a state-of-the-art Lilium™ eVTOL with 36 rotors, where the rotors are tilted around the front and rear wings. The manufacturer claims a range of 300 km and a speed of 300 km / h. However, this aircraft remains problematic because the high circular wing loading results in a low power-to-weight ratio (high installed power per unit weight), which reduces efficiency and range and generates a high noise level.

[0009] Instead of tilting the rotors around the wings, it is possible to arrange the rotors in a fixed position relative to the wings and tilt the wings. An example of this strategy is in Fig. Figure 6 is an artist's impression of an eight-rotor Airbus™ A3 Vahana. This aircraft is problematic because it trades higher efficiency in forward flight for a very high power requirement during the transition from vertical climb to forward flight. During such a transition, the wings act like giant airbrakes.

[0010] It is also possible to tilt the rotors around one or more fixed blades. Although no photo is available, it shows Fig. Figure 7 shows a computational fluid dynamics (CFD) solution for the Joby™ 6-rotor eVTOL concept aircraft. This aircraft solves some of the problems mentioned above, but the use of a multi-rotor strategy means that the rotors are relatively small. This inevitably results in a high circular area loading, leading to low power handling (high installed power per unit weight) and high noise levels.

[0011] The only other solution that the state of the art appears to have considered is the separation of the vertical lift rotors from the forward propulsion rotors / propellers. The idea behind this is that by using different lift and cruise propulsion systems, each system can be optimized for its specific function. Fig. Figure 8 is an artist's impression of the Aurora™ eVTOL concept aircraft, which uses eight hub rotors and a tail-facing propeller. This design is problematic because the dual propulsion systems require heavier and more expensive hardware, the climb rate in wing-assisted cruise is low due to the cruise engine's design for low-level cruising, and the stall speed in wing-assisted flight is potentially lower due to the design's optimization for a higher cruise lift coefficient (smaller wing). For example, at an airspeed of 130 mph, a vertical gust of 20 ft / s increases the angle of attack by 6 degrees and can stall a small wing at its efficient lift coefficient of 0.9, but not a larger wing at CL=0.5.

[0012] The Fig. 9A and Fig. Figure 9B shows artistic renderings of a similar design, the Terrafugia™ eVTOL aircraft. This design exhibits the aforementioned disadvantages regarding dual propulsion systems, and furthermore, the dual-tilt rotor configuration offers no method for angle-of-attack control during wing-assisted flight.

[0013] State-of-the-art motor installations are also designed for small rotors with low torque requirements. For example, [the text abruptly ends here, so the translation stops as well.] Fig. 10. The engine installation of the Airbus™ A3 Vahana™. The engine is arranged in a direct-drive configuration, where the engine and propeller rotate at the same speed. This simple drive system solution is problematic for large rotors with high torque requirements.

[0014] Due to the laws of physics, it is relatively easy to design a multi-rotor eVTOL aircraft capable of carrying a small payload (less than 500 pounds) over short distances. However, for larger payloads and commercially desirable ranges, the strategy of using many rotors becomes increasingly problematic. Using a larger number of smaller rotors provides a smaller rotor area than fewer large rotors, requires more power per unit of aircraft weight, makes low-noise hovering more difficult because the smaller overall rotor area necessitates higher Mach numbers at the blade tips, a greater number of wide-chord blades, or both, and makes autorotation after a power loss more dangerous, as the autorotation sink rate increases proportionally to the square root of the rotor area loading, and recovery from a high sink rate is risky.

[0015] Using a small number of larger rotors (two, three, or four) could solve some of the problems discussed above, but this approach is completely contrary to prevailing wisdom. Among other things, the characteristics required to optimize vertical lift differ significantly from those needed to optimize forward flight. Furthermore, experts would reject the idea of ​​a smaller number of rotors, arguing that it would unacceptably compromise safety in the event of a rotor engine failure and lead to unacceptable inefficiencies for an eVTOL aircraft. Moreover, these problems cannot be solved by separate lift and propulsion systems.

[0016] What is still needed is a vertical take-off and landing rotary-wing aircraft capable of safely carrying a payload of at least 500 pounds, using a unified lift and forward propulsion system with no more than four rotors, operating simultaneously with current battery technology.

[0017] All publications contained herein are incorporated by reference to the same extent as if each individual publication or patent application were expressly and individually indicated as incorporated by reference. If a definition or use of a term in an incorporated reference is inconsistent with or contradicts the definition of that term provided herein, the definition of that term provided herein shall prevail, and the definition of that term in the reference shall not apply.

[0018] In some embodiments, the numbers expressing quantities of components, properties such as concentration, reaction conditions, and so on, which are used to describe and claim certain embodiments of the invention, are to be understood as being modified in some cases by the term "approximately." Accordingly, in some embodiments, the numerical parameters specified in the written description and the appended claims are approximate values ​​that may vary depending on the desired properties to be achieved with a particular embodiment. In some embodiments, the numerical parameters should be interpreted taking into account the number of significant figures specified and applying standard rounding methods.Notwithstanding the fact that the numerical ranges and parameters representing the broad scope of some embodiments of the invention are approximate, the numerical values ​​presented in the specific examples are given as accurately as possible. The numerical values ​​presented in some embodiments of the invention may contain certain errors that inevitably result from the standard deviation observed in the respective test measurements.

[0019] As used in the description herein and in the following claims, the meaning of "ein", "eine" and "der", "die", "das" includes the plural unless the context clearly indicates otherwise. As used in the description herein, the meaning of "in" also includes "in" and "on" unless the context clearly indicates otherwise.

[0020] The mention of value ranges herein serves only as a shorthand notation to indicate each individual value falling within the range. Unless otherwise specified herein, each individual value shall be included in the specification as if it were individually listed herein. All procedures described herein may be carried out in any suitable order unless otherwise specified herein or clearly contradicted by the context. The use of examples or exemplary formulations (e.g., "as") relating to certain embodiments is for the sole purpose of better illustrating the invention and does not constitute a limitation of the scope of the otherwise claimed invention. No formulation in the description should be construed as indicating an unclaimed element that is essential to the exercise of the invention.Unless expressly stated otherwise, all regions include their endpoints, and open regions are to be interpreted as being limited at the open end by commercially feasible embodiments.

[0021] Groupings of alternative elements or embodiments of the invention disclosed herein are not to be construed as limitations. Each group member may be referenced and claimed individually or in any combination with other members of the group or other elements found herein. One or more members of a group may be included in or removed from a group for reasons of expediency and / or patentability. If such inclusion or deletion occurs, the specification shall be deemed to contain the group in the amended form and thus satisfy the written description of all Markush groups used in the appended claims.

[0022] As used herein, and unless the context specifies otherwise, the term "coupled to" is intended to include both direct coupling (where two coupled elements are touching each other) and indirect coupling (where at least one additional element is located between the two elements). Therefore, the terms "coupled to" and "coupled with" are used synonymously.

[0023] The following discussion provides many exemplary embodiments of the invention. Although each embodiment represents a single combination of inventive elements, the invention is considered to include all possible combinations of the disclosed elements. Thus, if one embodiment includes elements A, B, and C, and a second embodiment includes elements B and D, the invention is considered to include other remaining combinations of A, B, C, or D, even if these are not explicitly disclosed. Summary of the subject matter of the invention

[0024] The invention provides devices, systems and methods in which an electric, vertical take-off and landing (eVTOL) aircraft is designed to carry at least 500 pounds (approximately 227 kg) using a reduced number (e.g., 2-4) of variable-speed, rigid (non-articulated) rotors.

[0025] Various objects, features, aspects and advantages of the invention will become clearer from the following detailed description of preferred embodiments together with the accompanying drawings, in which the same numbers represent the same components. Brief description of the drawings Fig. 1A is a conceptual image of an emerging urban transportation market for the hybrid-electric, vertical take-off and landing (eVTOL) aircraft proposed by Uber®. Fig. 1B is a projected timetable for the development and operation of such an aircraft according to the state of the art. Fig. 1A. The Fig. 2A and Fig. 2B are artistic renderings of a state-of-the-art Volocopter™ with 16 rotors. Fig. Figure 3 is an artist's rendering of a state-of-the-art 8-rotor Ehang™. Fig. Figure 4 is an artist's rendering of a state-of-the-art CityAirbus™ with 8 rotors. Fig. Figure 5 is an artist's impression of a state-of-the-art 36-rotor Lilium™ eVTOL, in which the rotors tilt around the front and rear wings. Fig. Figure 6 is an artist's impression of a state-of-the-art 8-rotor Airbus™ A3 Vahana™. Fig. Figure 7 is a computational fluid dynamics (CFD) flow solution image for the state-of-the-art 6-rotor eVTOL concept aircraft Joby™. Fig. Figure 8 is an artist's impression of the Aurora™ eVTOL concept aircraft, which uses eight hub rotors and a rear-facing propeller. The Fig. 9A and Fig. Figure 9B shows artistic designs of a state-of-the-art Terrafugia™ eVTOL aircraft. Fig. Figure 10 shows the state-of-the-art engine installation of the Airbus™ A3 Vahana™. Fig. Figure 11 is a schematic perspective view of a preferred VTOL aircraft according to the inventive concepts presented herein. Fig. 12 is a table with the dimensions and parameters of the aircraft of Fig. 11. The Fig. 13A and Fig. 13B are schematic top and side views with dimensioned drawings of the aircraft by Fig. 11. Fig. 13C is a table that lists possible seating arrangements and calculated weights of the aircraft. Fig. 11 describes. The Fig. 13D and Fig. 13E are schematic perspective views of the aircraft by Fig. 11 with open doors and hatches. Fig. 13F is a schematic side view of an open rear ramp of the aircraft from Fig. 11. Fig. 13G is a schematic view of a side section of the aircraft of Fig. 11. Fig. Figure 14 is a schematic side view of a slotted flap, which is fitted to the aircraft of Fig. 11 can be used and is shown in four different positions. Fig. 15A is a schematic perspective view of an outer wing folding function, which is fitted to the aircraft of Fig. 11 can be used. Fig. 15B is a schematic front view of the outer wing of Fig. 15A in a folded orientation on the aircraft from Fig. 11. Fig. 15C is a schematic top view of the aircraft by Fig. 11, wherein the outer wing is in a folded orientation so that the aircraft fits on a circular parking space with a projected diameter of 45'. Fig. 16A is a table that shows the calculated geometry and properties of the rotor as functions of the dimensionless radial position of the aircraft. Fig. 11 describes. The Fig. 16B-16E are diagrams showing the calculated bending stiffness in the flap direction, the bending stiffness in the profile depth direction, the torsional stiffness, and the mass per unit length of the proposed rotor blades of the aircraft of Fig. 11. Fig. Figure 17 is a diagram showing the calculated lowest blade natural frequencies at a collective control setting of zero degrees as a function of the aircraft's rotor speed. Fig. 11 represents. Fig. Figure 18 shows schematic representations of five cross-sectional airfoil profiles of the rotor blades at specific radial positions of the aircraft. Fig. 11. Fig. Figure 19A is a first schematic perspective view of the propulsion system that is fitted to the aircraft by Fig. 11 can be used, which is housed in a streamlined gondola in a position required for wing-supported cruise flight. Fig. 19B is a second schematic perspective view of a part of the propulsion system that is attached to the aircraft by Fig. 11 can be used. Fig. Figure 20 is a diagram showing the calculated effects of engine speed on engine weight for the aircraft of Fig. 11 represents. Fig. 21A is a schematic representation of a vertical cross-section of a preferred configuration with individual blade control (IBC), which is used with the aircraft of Fig. 11 can be used. Fig. 21b is a schematic representation of a vertical cross-section of an alternative, preferred single-blade control (IBC) configuration, which is connected to the aircraft of Fig. 11 can be used. Fig. Figure 22 is a schematic perspective view of the aircraft of Fig. 11 with actuators for single-sheet control (IBC) according to the Fig. 21A or Fig. 21B with four-bladed primary and four-bladed secondary rotors. Fig. 23A is a schematic cross-sectional view of a gondola attached to the aircraft of Fig. 11 can be used, in which a battery is arranged below the wing and inside the gondola. Fig. 23B is a schematic cross-sectional view of a gondola and wing attached to the aircraft of Fig. 11 can be used, in which a battery is arranged inside the wing. The Fig. Figures 24A-24G are schematic perspective views of an alternative, preferred VTOL aircraft according to the inventive concepts presented herein. This version does not have secondary rotors. The Fig. 25A and Fig. Figure 25B are schematic vertical cross-sectional views of the alternative VTOL aircraft of the Fig. 24A-24G, which represent a 3-row seating arrangement, compatible with the 4-rotor configuration of Fig. 13G is comparable. Fig. 26 is a table with the dimensions and parameters of the alternative VTOL aircraft with 2 rotors of the Fig. 24A-24G. Detailed description

[0026] The invention provides devices, systems, and methods in which an electric vertical takeoff and landing (eVTOL) aircraft is designed to carry at least 500 pounds (approximately 227 kg) using a reduced number (2-4) of variable-speed, rigid (non-articulated) rotors, generally mounted as primary and secondary rotors. The rotors, whether primary or secondary, are preferably tiltable rotors, such that one or more of the rotors can provide a significant portion of the lift (e.g., 70%, etc.) during rotor-assisted flight (e.g., vertical takeoff, etc.) and can be tilted to provide forward thrust (or air braking) during wing-assisted flight.

[0027] In some considered embodiments, each rotor can be driven by its own electric motor or motors, and in other considered embodiments, several rotors can be driven by a single electric motor. In particularly preferred embodiments, the individual rotors can be driven by three electric motors. It is also considered that different electric motors may be powered by different battery sets, or that several electric motors may be powered by a single battery set.

[0028] The terms "battery" and "battery pack" are used interchangeably herein and refer to one or more chemical cells that generate electricity. The batteries preferably use lithium-ion chemistries and have a specific energy density of about 100 kWh / lb. Other battery chemistries under consideration include lithium polymer and lithium metal.

[0029] Non-articulated rotors are preferred because the variation in individual blade angles can be used to generate torque and thus control the aircraft's pitch in both VTOL and wing-based cruise flight. Blade angle control is preferably achieved by individual blade control actuators, which are preferably integrated into the respective blades and mounted axially to the pitch axis. The individual blade control system, employed on at least each of the first and second primary rotors, imparts a differentiated collective angle of attack to the blades on the rotor, thus maintaining a nearly constant rotor thrust while increasing the shaft torque above the torque required without a differentiated collective angle of attack.Details can be found in the pending preliminary applications 62 / 513930 (Tigner) “A Propeller Or Rotor In Axial Flight For The Purpose Of Aerodynamic Braking” and 62 / 513925 (Tigner) “Use Of Individual Blade Control To Enhance Rotorcraft Power Response Quickness”, which are incorporated herein by reference in their entirety.

[0030] In preferred embodiments, the aircraft has a primary and a secondary rotor. The primary rotors comprise blades and hubs designed to provide torques that are at least equal to the maximum lift of the rotor multiplied by 6% of the rotor radius, more preferably at least 9% of the rotor radius, and most preferably at least 12% of the rotor radius.

[0031] To achieve commercially viable flight duration, lift, and other characteristics with no more than four rotors and currently available battery technologies, at least the primary rotors must be relatively large. Accordingly, each primary rotor is designed to provide a circular wing loading of less than 10 psf and a hovering power load of more than 8 lb / HP. More preferably, each primary rotor is designed to provide a circular wing loading of less than 6 psf and a hovering power load of more than 10 lb / HP. Other considered embodiments of the aircraft have a power load of less than 8 lb / HP.

[0032] To achieve high rotor efficiency in rotor-assisted and wing-assisted flight, continuous rotor operation over a wide rotor speed range (such as 20% to 100%) is also required, with the considered embodiments employing rotor designs disclosed in US patents 6007298 (Karem) “Optimum Speed ​​Rotor” (OSR) and 6641365 (Karem) “Optimum Speed ​​Tilt Rotor” (OSTR).

[0033] Applying the teachings of OSR and OSTR, the aircraft considered herein preferably achieve an impact stiffness of each blade that is not less than the product of 100 times, or more preferably 200 times, the rotor diameter in feet raised to the fourth power, measured in pounds per square inch, at 30% of the rotor radius, measured from a center of rotation of the rotor.

[0034] Applying the teachings on OSR and OSTR, the weight of each blade in pounds preferably does not exceed the product of 0.004 times the diameter of the rotor in cubic feet.

[0035] Embodiments comprising first and second primary rotors are considered to include at least one optional first auxiliary rotor, each of which has no more than 50% of the circular area of ​​each of the primary rotors. In preferred embodiments, each of the auxiliary rotors has no more than 40% of the circular area of ​​each of the primary rotors. The auxiliary rotors need not be of the same size.

[0036] The auxiliary rotor(s) is / are also preferably rigid (non-articulated) rotors designed to generate pitching moments by changing the angle of attack of individual blades. At least the first auxiliary rotor is advantageously designed to provide a maximum aircraft pitching moment that is not greater than the collective total aircraft pitching moment capacity of the primary rotors.

[0037] For each rotor, the rotating hub, the corresponding hub bearing, the gearbox, and the motor mounting device are designed together as an integrated rotor drive system. The preferred embodiment includes three independently controlled motors connected to a single gearbox per primary rotor. The three independently controlled motors offer a safety advantage through redundancy, and furthermore, this configuration has proven to be a lightweight solution for the high-torque output required for a variable-speed rotor.

[0038] Preferred embodiments include a wing supporting at least one first and one second rotor, each of which is arranged in a rotor assembly designed to tilt at least 90° relative to the wing. In particularly preferred embodiments, corresponding motors or other propulsion systems are designed to tilt together with the rotor assemblies. At least the primary rotors are open, meaning that when the rotors tilt, they are not circumferentially constrained by an air guide.

[0039] As with the rotors, the wing is also relatively large in relation to the aircraft's weight and payload. For example, it is preferred that the wing be dimensioned and designed to provide a wing loading of no more than 40 psf and a wing-assisted stall speed of no more than 90 KIAS. Particularly preferred wings are further designed to provide a wing loading of no more than 20 psf and a wing-assisted stall speed of no more than 50 KIAS. Preferred wings are further designed to provide an airspeed range of no less than 20 KIAS during the transition from fully rotor-assisted straight and level flight to fully wing-assisted straight and level flight, and a lift-to-drag ratio of no less than 10 during wing-assisted cruise flight. Particularly preferred wings are further designed to provide a transition airspeed range of no less than 40 KIAS.

[0040] To further reduce the aircraft's stall speed, the preferred wing is equipped with an actuated slotted flap. In the particularly preferred configuration, the flap can be used to provide roll control for the aircraft. The wing is preferably designed with wingtip sections incorporating a control system and electrical or other actuators that adjust the wingtips to a negative dihedral angle between 20 and 90 degrees to: (a) provide reduced wing download during hovering; (b) provide roll assistance when taxiing in crosswinds; and (c) provide aircraft anchorage.

[0041] The wing, rotors, and other components and features discussed herein are preferably designed such that the aircraft can maneuver at 3 g at maximum weight without loss of altitude or speed, yet still provide a low sustained autorotation sink rate in the event of an engine failure. The preferred embodiment has a sustained autorotation sink rate of less than 1,000 ft / min.

[0042] In some embodiments, at least one first battery or other energy source is arranged in the wing. In some embodiments, a landing gear extends from at least one of the fuselage and the wing.

[0043] In another preferred embodiment, at least one first battery or other energy source is arranged in a primary rotor nacelle.

[0044] Embodiments are considered which include a tail assembly and / or a canard, each preferably having a lift surface with an area between 10% and 100% of that of the wing.

[0045] The embodiments under consideration include both manned and unmanned aircraft. If a fuselage is present, it can therefore include a passenger compartment with at least one seat designed to provide seating for a person.

[0046] Electronic controls are also being considered that would be sufficient to fly the aircraft without a human pilot on board.

[0047] Fig. Figure 11 is a perspective view of a preferred VTOL aircraft according to the inventive concepts presented herein. The aircraft comprises a wing 1101, a static nacelle 1102, a tilting nacelle 1103, a fuselage 1150, a tail surface 1130, and a first tilting rotor system 1110. A particularly preferred embodiment includes a first tilting auxiliary rotor system 1140.

[0048] The rotor system 1110 includes rotor blades 1120. The rotor blades are of a rigid, hingeless design, including, for example, the design described in US Patent No. 6641365 (Karem). The rotor system provides collective thrust, as indicated by arrow 1113 and force moment 1114. The moments and forces can be controlled by rotating the blades about a spring axis 1121 extending along the length of the blade 1120. The angle of attack about the spring axis 1121 is represented by arrow 1122. The tip of the rotor blade follows a rotational trajectory represented by a circle 1116. The rotor blades 1120 and the tiltable nacelle 1103 can tilt about the tilt axis 1112 along the path shown by arrow 1111. To illustrate the function of the tilt rotor, the right gondola is in a wing-assisted flight orientation, while the left gondola is in a rotor-assisted flight orientation.The gondolas would be in similar orientations during normal operation.

[0049] The wing 1101 transfers the loads of the rotor system to the fuselage 1150. The fuselage 1150 is designed to carry payload and passengers and contains various systems, including landing gear.

[0050] Fig. Figure 12 is a table containing the dimensions and parameters of a preferred embodiment, where “*” denotes turbulent flow. The preferred embodiment described in the table is designed for a nominal payload of approximately 1,100 pounds and a basic launch weight of 4,767 pounds. A wing area of ​​250 square feet results in a wing loading of 19.1 psf. With a total wing area of ​​849 square feet, the hovering wing loading is 6.62 lb / ft. 2 , when one considers the effects of rotor wash on the airframe.

[0051] The Fig. 13A and Fig. Figure 13B shows a top view and a side view with dimensioned drawings of the preferred embodiment, which is shown in Fig. The aircraft depicted in Figure 11 is consistent. 1310 is a main landing gear wheel attached to fuselage 1150. 1311 is a nose landing gear wheel attached to fuselage 1150. All previously numbered elements are as described above.

[0052] Fig. Figure 13B is a side-view section showing the two upward-opening nose doors 1312, the four car-like doors 1313 and 1314, and the baggage compartment 1315 of the preferred embodiment of the aircraft with three possible configurations: a) air taxi with one pilot and four passengers, b) family use with up to eight passengers, and c) cargo / medical transport use with folding seats in the rear row and an optional ramp replacing the baggage compartment. The family-use configuration, with a payload capacity of 1,350 pounds, is intended to accommodate a family much like a large SUV, except that the aircraft, with its two nose doors, provides the equivalent of six doors compared to the SUV's four.

[0053] Fig. 13C is a table that describes the preferred seating arrangement of the aircraft and the calculated weights.

[0054] The payload of 1,350 pounds, the desired loading flexibility for the three rows of seats, and especially the rear loading of the baggage compartment or tail ramp with 400 pounds, result in a large shift of the aircraft's center of gravity (load vector) of up to 8.5 inches (13.2% of the mean aerodynamic chord). Maintaining the aircraft's stability and control during such a large shift of the center of gravity is achieved in rotor-assisted flight by the powerful combination of the angle-of-attack control of the preferred embodiment of the auxiliary rotors and the pitching moment of the rigid primary rotors, and in wing-assisted flight by the powerful combination of the angle-of-attack control of the large elevators and the pitching moment of the rigid primary rotors.

[0055] The Fig. 13D and Fig. Figure 13E shows views of the preferred embodiment of the aircraft with the doors and hatches open.

[0056] Fig. Figure 13F shows the rear ramp 1316 in the open position. The nose landing gear 1311 can optionally be adjusted in height, allowing the angle between the fuselage and the ground to be adjusted, thus providing additional clearance at the ramp opening.

[0057] Fig. Figure 13G shows a side-section view of the fuselage with a proposed seating arrangement, which is equipped with Fig. 13C is consistent. Electronics 1317, which is capable of flying the aircraft without a human pilot on board, is intended for future operation. Aerodynamic design

[0058] The aircraft considered herein are designed for efficient vertical and cruise flight. Furthermore, such aircraft are designed to ensure safe flight and to perform well in the intermediate state known as the "transition" between fully wing-based and rotor-based flight.

[0059] The rotor thrust required for vertical flight is on the order of ten times that required for efficient cruise flight. The preferred embodiment of the aircraft uses the variable-speed rotor described in US Patent 6,641,365 (Karem) to achieve high efficiency from 100 rpm in slow, wing-supported flight up to 460 rpm in hovering at 12,000 feet. The aerodynamic design of the rotor represents a relatively small trade-off between the optimal hovering and cruise characteristics, typical for the 5:1 speed ratio available with such a rotor. This requires a combination of airfoil designs with linear lift characteristics over a wide angle-of-attack and twist range, as well as chord distributions that balance the requirements for vertical and cruise flight.Tools for the design and analysis of airfoil sections, such as XFOIL, can be used to design and investigate airfoils that meet the desired properties. Software for analyzing rotors for cruise flight, such as XROTOR, and software for rotor performance in hovering flight, such as CHARM (CDI), can be used to optimize the rotor geometry for the desired performance characteristics. The resulting preferred rotor geometry is described in [reference missing]. Fig. 16A as a table and the wing sections in Fig. 18 shown.

[0060] For efficient cruising flight, a high lift-to-drag ratio of at least 10 is desirable. The drag of the fuselage and nacelle is minimized using computational fluid dynamics (CFD) software, such as STAR-CCM+, to analyze and iteratively optimize the shape, taking into account practical considerations like volume for propulsion and payload, as well as structural requirements. The airfoil can be optimized using airfoil tools like the aforementioned XFOIL. Considerations for wing optimization include a trade-off between drag during cruising flight, downward load during vertical flight, maximum lift during transitions, and structural requirements.

[0061] A preferred method for increasing maximum lift during transition without negatively affecting drag during cruise flight is a slotted flap, as shown in the sectional drawing in Fig. Figure 14 shows several flap deflection angle positions: maximum upward deflection (-8°), no deflection (0°), maximum lift coefficient (+22°), and maximum downward deflection (+65°). The flap 1401 rotates about a simple hinge 1404. In the retracted position (position 0), the flap 1401 causes only minimal additional drag compared to a one-piece airfoil. When deployed at an optimal high lift angle (position +22°), a gap is exposed through which the airflow from the underside of the first element 1402 can flow over the flap element 1401. In the preferred embodiment, the shape of the flap and the gap are designed to provide a linear lift response to small deflection angles, so that the slotted flaps can also be precisely controlled as ailerons for roll control of the aircraft.The preferred slotted flap extends spanwise from the fuselage side to the wingtip, with a break at the engine nacelle. The flap is divided into several spanwise sections to reduce stresses caused by wing deformation. The flap hinge is set back from the wing surface such that a gap opens when the flap is deflected downwards. An elastic surface seal 1403 minimizes drag when the flap is in its retracted position.

[0062] The most critical phase for flight safety in an eVTOL flight with wings is the transition from fully rotor-assisted to wing-assisted flight at a safe forward speed. This is particularly important in turbulent, windy conditions in urban environments at low altitudes. Unlike previous technologies, the use of large primary rotors and a large wing area together ensure a safe transition.

[0063] In the preferred embodiment (with two auxiliary rotors), the use of large rotors results in a circular area loading that enables a low noise level (350 rpm, Mach number at the rotor tip below 0.35), efficient, stable hovering, and rotor-assisted maneuvering at 2 g at 495 rpm. This is combined with a large wing (250 ft). 2With a slotted flap at 22 degrees, the aircraft is designed to achieve a stall speed of 50 KIAS at a takeoff weight of 4,767 pounds. At a rotor speed of 550 rpm, the aircraft can be fully rotor-supported (no wing or empennage lift) at 90 KIAS, 40 KIAS higher than the minimum wing-supported speed, and can sustain an instantaneous rotor lift of 2.5 g. At 90 KIAS, the aircraft can exhibit a wing-supported lift of 3.25 g. These large margins avoid most low-lift accidents and control accidents typical of low-speed flight and transitions in turbulent weather. Folding outer wing

[0064] The outer wing folding function is in Fig. Figure 15A shows the outer wing 1501 being foldable about the hinge line 1503 relative to the inner wing 1502, and the movement is controlled by a folding actuator (not shown). The folding wing drive is designed to withstand flight and ground loads. It has a spring-loaded skid 1504 at the wingtip that makes contact with the ground during landing. The large wingspan, required for efficient flight and acceptable transition characteristics, makes the aircraft susceptible to crosswinds and ground gusts. Skids at the tips provide the aircraft with additional ground stability and safety. The wingtip also incorporates a mooring element (not shown) to secure the aircraft during parking. In rotor-assisted flight, the downward load due to rotor downwash on the outer wings is reduced, since the downward load is inversely proportional to the separation distance between the rotor and the wing. Fig. Figure 15B shows the front view of the aircraft with the outer wings folded. Furthermore, the folded aircraft fits onto a smaller footprint. Fig. Figure 15C shows the top view of the aircraft, which fits into a circular parking space with a projected diameter of 45'. Leaf construction

[0065] Fig. Figure 16A shows the geometry and properties of the rotor as a function of the dimensionless radial position. Fig. Figure 18 shows cross-sectional wing profiles of the rotor blade at certain radial positions.

[0066] Fig. 16B represents the bending stiffness in the direction of impact, or perpendicular to the profile depth, of the exemplary embodiment of the rotor blade from the radial position zero at the root to the radial position 1 at the tip. Fig. 16C represents the pivoting bending stiffness, or in the profile depth direction, of the exemplary embodiment of the rotor blade from the radial position zero at the root to the radial position 1 at the tip. Fig. Figure 16D represents the torsional stiffness of the exemplary embodiment of the rotor blade from the radial position zero at the root to the radial position 1 at the tip. Fig. 16E represents the mass per unit length of the exemplary embodiment of the rotor blade from radial position zero at the root to radial position 1 at the tip. Fig. 16F represents the gravity position in the profile depth direction relative to the blade pitch axis of the exemplary embodiment of the rotor blade from radial position zero at the root to radial position 1 at the tip.

[0067] The Fig. Figures 16C-E represent the bending stiffness in the flap direction, the bending stiffness in the profile depth direction, and the torsional stiffness of the proposed blade embodiments. High stiffness relative to mass is required to avoid structural dynamic problems during rotor operation over a wide speed range. The blade mass distribution of the proposed blade embodiment is shown in Fig. Figure 16F shows that, to avoid aeroelastic instabilities, the center of gravity of the rotor blade must not be located much further aft in the chord direction than the blade pitch or spring axis. It was found that the rotor blade design is free of aeroelastic instabilities under operating conditions, with the center of gravity as shown in Figure 16F. Fig. 16F is shown as balanced.

[0068] Fig. Figure 17 shows the lowest blade natural frequencies at a collective setting of zero as a function of rotor speed. The rays emanating from the origin represent the rotor's harmonic frequencies 1 / U, 2 / U ... 10 / U. The operating speed ranges are marked on the horizontal axis. The stiffness and mass distributions of the Fig. The 16A-F series has the greatest influence on the natural frequencies of the rotor blades. The natural frequencies remain well separated from each other across the entire operating range. Due to the design with high stiffness relative to weight, the natural frequencies are significantly higher than those of typical rotor blades. The first flapping mode remains above the rotor's excitation frequency of 3 / U across the entire operating range, while the less damped first floating mode remains above the excitation frequency of 4 / U. This separation above the primary rotor's excitation frequency of 3 / U allows operation over a wide rotor speed range without excessive vibration loads or vibrations due to resonance.

[0069] Software programs for simulating and optimizing rotor dynamics, such as CHARM and CAMRAD, can be used to iterate the rotor blade design, taking into account the desired described properties. Finite element analysis (FEA) software can be used for more accurate structural analysis, and CFD codes can be used for more accurate aerodynamic analysis and refinement.

[0070] The preferred auxiliary rotor and blades are designed to the same performance specifications as the primary rotor, but with a smaller diameter. Hub drive system

[0071] The propulsion system is housed in a streamlined gondola, which is located in Fig. Figure 19A illustrates the position required for wing-assisted flight. The rotor's axis of rotation is shown as XX and the direction of flight by arrow A. For aircraft with tilting rotors, the rotor's thrust axis must, by definition, be rotated from the horizontal flight position to the vertical lift position. This angle is not less than 90 degrees and can be 105 degrees or more. The axis around which the forward section of the nacelle, including all propulsion components, tilts is shown as YY.

[0072] The blade arms 1901 are mounted in spring-bearing mounting brackets 1902, which are bolted to the rotating hub 1903, which rests on a large-diameter bearing 1904. The three motors 1905 are arranged symmetrically around the hub's center, one of which, 1906, is shown in section. The output sun gear 1907 is driven via a freewheel clutch 1908. The planet gears 1909 are mounted in a planet carrier 1910, which is attached to the output pinion 1911. The three identical output pinions engage with the ring gear 1912. The hub loads are transferred from the hub bearing via the intermediate structure 1913, which is attached to the nacelle 1914, made of a monocoque composite construction, by bonding and riveting. The shell structure is attached to the rear gondola at the hinge points 1915, with the tilt drive frame 1916 connecting both gondola elements at the drive mounting bracket 1917.The electronic motor control boxes 1918 are individually packaged for redundancy and have phase connections 1919 to the motors. The connection for the motor's liquid cooling 1920 is illustrated, as is the oil drip tray 1921. The alternative tilt actuator 1922 is shown mounted on a transverse rotary axis.

[0073] Fig. Figure 19B presents an alternative perspective view of a section of the in Fig. 19A shows the streamlined gondola. Fig. Figure 19B provides a close-up view of the leaf legs 1901 as mounted in the spring bearing retaining bracket 1902, and of the screw connection of the spring bearing retaining bracket 1902 to the rotating hub 1903, which rests on a large diameter bearing 1904.

[0074] The entire rotor hub, including the blade spring bearings and the pitch-angle drive system coupled to the electric drive, forms an integrated assembly. The system is illustrated as a three-blade arrangement; other numbers of blades are installed similarly. The four predominant loads distributed through the assembly from the bogie to the nacelle structure are the blade impact loads, the mast moment, the thrust or lift vector, and the drive torque. A large-diameter moment bearing connects the rotating hub elements to the nacelle structure. The large, slow-rotating rotor creates a drive condition in which the rotor's torque / speed characteristic far exceeds the capabilities of a direct-drive motor. A summary of the analysis of the effects of motor speed on motor weight is provided in Fig. Figure 20 shows that, at constant power, the high-speed motor with gear reduction offers a weight reduction with increasing gear ratio, and further weight reduction is achieved by multiplexing the motors. Since flight safety is of paramount importance, multiplex motors offer complete electrical redundancy. A degree of mechanical redundancy is also achieved by installing a one-way clutch (freewheel clutch) on each of the motor output shafts. The illustrated system features three motors but is adaptable to a larger number of motors.

[0075] The typical application of a direct-drive electric motor for eVTOL lift rotors offers several advantages. It is simple, and with simplicity comes inherent reliability. Furthermore, the added weight of a gearbox is avoided. However, weight reduction is also possible by increasing the electric motor's speed and by using a geared output. By trading torque for speed at constant power, significant weight savings are achievable. Weight reduction is limited by the reduced cooling options as the motor size decreases. Therefore, the choice between a direct drive and a geared motor for weight optimization depends on the desired power output and speed. At lower output speeds, a geared drive offers a weight advantage, while at higher speeds, a direct drive is more weight-efficient. Fig. Figure 20 shows the weight of the motors and drive system for the 300 HP rated power of the primary rotor in the preferred embodiment. In this case, the weight limit is approximately 2000 rpm. In the preferred embodiment, the primary rotor speed is 400-460 rpm in high-performance hovering and 350 rpm in high-performance wing-assisted climb, which clearly favors a geared rotor drive.

[0076] There are several advantages to incorporating more than one motor into the design. A) More motors can be more weight-efficient. The ability to dissipate heat depends on the motor's surface area, and the power output at constant speed depends on the motor's volume and therefore its weight. A larger surface-to-volume ratio allows for better cooling. If the minimum weight is primarily determined by cooling capacity, more motors offer better weight efficiency because they have a better surface-to-volume ratio. B) Reliability can be increased through redundancy. In a configuration where acceptable output power can be maintained if one or more drive motors fail, overall reliability is increased by redundancy. However, the complexity of many motors can reduce reliability.

[0077] The higher the gear ratio, the greater the weight saving. Since the weight of the gearbox is driven by the high-torque output shaft, the gearbox weight is primarily independent of the gear ratio. As in Fig. As shown in Figure 20, the weight of the motors is inversely proportional to the rotational speed. However, the weight savings are limited by the reduced heat dissipation capacity as the size decreases. Furthermore, there are practical limitations regarding motor speed, including: keeping the magnets under high centrifugal force (in motors of this type), limitations on available bearing speeds, and limitations on the electronic switching speeds for motor commutation.

[0078] In the present example, a gear ratio of 20:1 reduces weight while simultaneously mitigating the challenges of very high engine speeds. In other considered embodiments, the transmission could have gear ratios of 3:1, 5:1, 10:1, 20:1, or 30:1.

[0079] With an overall gear ratio of nearly or greater than 20:1, two reduction stages are required to fully exploit the weight savings of high-speed motors. Each motor is equipped with a planetary reduction gear set that drives a combination ring gear mounted on the hub. All features of the assembly are optimized for minimal weight; for example, the use of three drive pinions engaging with a single large ring gear minimizes the ring gear's surface area, resulting in material savings.

[0080] The motors, their control electronics, and the transmission all require cooling. The preferred fluid for motor and electronics cooling is water / glycol, and a separate liquid-to-oil heat exchanger is used for transmission oil cooling. The transmission oil is located in an oil pan at the lower rear of the transmission housing.

[0081] The gondola tilting system is shown as a system of three linear actuators, with the rear pair providing 60 degrees of gondola deflection and the front actuator providing the remaining 55 degrees. An alternative system is the use of a high-torque rotary actuator operating via a four-bar linkage. The rotary actuator, US7871033 (Karem et al.), which describes its implementation in detail, is cited in the references. Single-sheet control

[0082] In the preferred embodiment, the actuators for the individual blade control (IBC) 2101 enable precise, independent control of the rotor blade paths. The rotor torque and forces can be controlled by independently controlling the blade angle. Fig. Figure 22 shows a similar image to the one in Fig. Figure 11 is shown, with the exception that the first tilt rotor system 2210 and the first tilt auxiliary rotor system 2240 are 4-blade rotor applications. The application of the IBC drive with a 4-blade rotor enables the deceleration of the aircraft (negative rotor thrust) in wing-assisted flight without generating large hub moments. All elements similar to those in Fig. The applications numbered 11 are as described above. Details can be found in the pending preliminary applications 62 / 513930 (Tigner) “A Propeller Or Rotor In Axial Flight For The Purpose Of Aerodynamic Braking” and 62 / 513925 (Tigner) “Use Of Individual Blade Control To Enhance Rotorcraft Power Response Quickness”, which are incorporated herein in their entirety by reference.

[0083] A preferred IBC configuration is in Fig. Figure 21A shows the design approach of integrating an electric actuator into the blade itself, positioned so that the actuator and leaf spring axes coincide. Certain blade construction requirements must be met for this approach to be feasible. OSTR rotor blades should, to their great advantage, exhibit high stiffness in flapping and pivoting, resulting in much greater chord and thickness in the blade root section compared to conventional rotor blades. The resulting blade spar is hollow and has a suitable diameter to comfortably accommodate the cylindrical electric actuator. Combined with a reduction gearbox, the electric motor drive can connect the blade to the hub in the direction of rotation without any mechanical links and can be commanded and controlled just like other flight control actuators.The general term for this type of propulsion is single-blade control (IBC), which enables a completely new and optimized matrix of blade azimuth and angle of attack. This results in aerodynamic advantages.

[0084] Furthermore, military helicopters intended for ship-based operations must be designed to be compact by means of folding. Folding an existing rotor blade according to the prior art, while maintaining the integrity of the pitch linkage, results in a complex arrangement of mechanical parts. The present invention eliminates this complexity, and the only new requirement for the design of the actuator located inside the blade is that the electrical wiring bends with the folding angle. This requirement can be met easily and simply.

[0085] The hollow blade spar 2101 is inserted into the receiving bore of the hub 2102. It rests on the blade via the bearing 2103, which runs on the inner ring 2104, and is sealed by the gasket 2105. When the blade needs to be folded, the blade and the hub area rotate around the hinge 2106. Either one motor stator 2107 or two motor stators 2107 and 2108 drive the rotor 2109, which is guided by the tail bearing 2110 and the rotor bearing 2111. The position of the motor rotor, and thus the angular position of the blade, is detected by the position sensor 2112 using a static reference via the stationary core 2113. The motors drive the gearbox 2114, which is secured to the blade root by the fixings 2115. The reaction torque of the gearbox is absorbed by an elastic coupling 2116, the purpose of which is to isolate the gearbox from torque-induced deflections caused by impact and pivoting loads of the blade.Centrifugal loads and torque-induced radial loads are absorbed by tapered roller bearings 2117. The blade's drive torque is absorbed via the toothing 2117, and the centrifugal force is absorbed via the nut 2119. The elastic electrical connection cable 2120 transmits the motor power and control information from the slip ring 2121, which rotates about the hub rotation axis 2122 (shown by axis XX), with the static part of the slip ring resting on the airframe structure 2123. The supply and return lines for the coolant 2124 are routed through the rotary feedthrough 2125.

[0086] Fig. Figure 21B shows an alternative component arrangement. If the rotor blade does not need to be folded and the blade spring axis is precisely controlled relative to the hub by rigid spring bearings, the pitch-angle actuator can be mounted on the hub instead of the blade. In this arrangement, the actuator assembly, consisting of the motor(s), reduction gear, and the necessary sensors and connecting cables, is attached to the hub. This connection is torsionally rigid but flexible in its orientation to accommodate the deflections that occur with highly stressed blades. The toothed output pulley is connected to a toothed section inside the blade. There are practical system advantages to mounting several actuators together on a single hub, as they can be electrically interconnected via common driver, power, and cooling paths.

[0087] The cylindrical blade spar 2131 rests on the outer spring bearing assembly, which consists of the bearing retainer bracket 2132, the outer ring 2133, the rollers and cage 2134, the inner ring 2135, and the seals 2136. The blade root is stabilized by an inner diaphragm 2137, which is secured with rivets 2138. The diaphragm has internal teeth at 2139 for transmitting torque from the elastic drive bellows 2140. This is the separation point when the removed blade is retracted over the stationary actuator.

[0088] The split leaf retaining clip 2141 secures the inner root screw 2142 to the outer ring of the inner spring bearing 2143. The tapered roller and the cage 2144 run on the inner ring 2145, which is sealed by the seal 2146. The bearing preload is provided by the disc spring 2147, which acts on the thrust washer 2148.

[0089] The centrifugal force of the blade is absorbed by the actuator housing 2149, which is held by the mounting set 2150. This mounting set also secures the static core 2151 to the rotating hub component 2152. The static core carries both the motor stator windings 2153 and the position sensor 2154. The electric motor rotor 2155 rests on a journal bearing 2156 and a tail bearing 2157 and drives the reduction gear 2158.

[0090] A tubular extension 2159, which is non-rotatably attached to the airframe structure, carries a slip ring 2160 that supplies power and control signals to the actuator via a fixed cable harness 2161. battery

[0091] The preferred battery installation is shown in the cross-sectional view of the gondola in Fig. 23A shows the battery 2301 located below the wing 2302 and inside the nacelle 2303. The direction of flight is indicated by the block arrow B. The rear volume of the nacelle contains sufficient space to accommodate parts of the cooling system required for the electric propulsion system. Alternatively, the battery 2311 can also be contained within the wing structure 2312, as shown in Fig. 23B is shown. Battery 2311 has smaller cross-sectional dimensions than in Fig. 23A, but offers the same volume because it is enclosed by the long wing 2321. The nacelle 2313 can then be significantly smaller and have lower drag than in the primary preferred embodiment. Alternatively, a nacelle arrangement with an internal combustion engine and a generator (hybrid propulsion) can be used to power the electric motors and other aircraft systems, providing a significantly greater range. Alternative configuration

[0092] The Fig. References 24A-G disclose an alternative preferred embodiment without auxiliary rotors. This preferred embodiment utilizes the powerful angle-of-attack control of the primary rotors 2401 in rotor-assisted flight and the angle-of-attack control of a long-arm canard wing 2411 in wing-assisted flight to enable an aircraft with lower weight, lower drag, lower installed power, and lower cost. It offers the same aircraft performance but less control over the aircraft during transitions and wind gusts, as well as less adaptability to center of gravity shifts and payload variety (no cargo ramp and reduced baggage volume and weight) compared to the preferred embodiment with the auxiliary rotors and the large tail surface.

[0093] The alternative preferred embodiment consists of a fuselage 2421, a wing 2431, rotor blades 2401, and a canard 2411. The internal configuration is similar to the primary preferred embodiment. The fuselage has three rows of seats: a front row 2501, a middle row 2502, and a rear row 2503.

[0094] While the cabin volume of the alternative 2-rotor configuration is comparable to that of the 4-rotor configuration, wing-based drag is reduced by: a) a reduction in wing area of ​​250 ft² 2 at 140 ft 2, b) the absence of a tail assembly area, c) the absence of auxiliary rotor nacelles, d) the attachment of the wings to the fuselage behind the cabin (smaller frontal area), e) the possibility of extensive laminar flow over the fuselage, and f) lower drag in cruise flight due to the lower cruise weight (estimated 817 pounds lighter due to a lighter airframe and a smaller battery).

[0095] The Fig. 25A and Fig. Figure 25B shows the internal profile of the alternative 2-rotor configuration, which has a 3-row seating arrangement comparable to that of the 4-rotor configuration. Fig. 26 is a table with the dimensions and parameters of the alternative configuration with 2 rotors. Modifications

[0096] It should be clear to experts that, in addition to the modifications already described, many more are possible without deviating from the inventive concepts contained herein. The subject matter of the invention is therefore not to be limited, except as defined in the attached claims. QUOTES INCLUDED IN THE DESCRIPTION

[0000] This list of documents cited by the applicant was automatically generated and is included solely for the reader's convenience. The list is not part of the German patent or utility model application. The DPMA accepts no liability for any errors or omissions. Cited patent literature

[0000] US 62 / 509666

[0001] US 62 / 509674

[0001] US 62 / 656971

[0001] US 6007298

[0032] US 6641365 [0048, 0059] US 7871033

[0081] WO 62 / 513930

[0082]

Claims

[1] A VTOL aircraft, comprising: at least one wing mechanically coupled to a fuselage; at least four variable-speed tilt rotors, together designed and dimensioned to support a payload of at least 500 pounds; at least one first and at least one second motor, wherein the first and / or second motor is configured to drive a first of the at least four variable-speed tilting rotors; and at least one third and at least one fourth motor, wherein the third and / or fourth motor is configured to drive a second of the at least four variable-speed tilting rotors, wherein in a rotor-supported flight the at least four variable-speed tilt rotors are arranged together to lift the VTOL aircraft, wherein three of the at least four variable-speed tilt rotors are capable of providing at least 70% of the lift required to lift the VTOL aircraft carrying a payload of at least 500 pounds. [2] The VTOL aircraft according to claim 1, wherein at least one of the first, second, third and fourth engines comprises an electric motor. [3] The VTOL aircraft according to claim 2, wherein the VTOL aircraft further comprises a generator configured to supply power to the electric motor during flight. [4] The VTOL aircraft according to claim 2, wherein the first, second, third and fourth motors comprise an electric motor. [5] The VTOL aircraft according to claim 1, wherein the VTOL aircraft further comprises an internal combustion engine configured to supply power to at least one of the first, second, third and fourth engines during flight. [6] The VTOL aircraft according to claim 1, wherein the VTOL aircraft comprises an electrically powered VTOL aircraft. [7] The VTOL aircraft according to claim 6, wherein the electrically powered VTOL aircraft comprises a hybrid-electrically powered VTOL aircraft. [8] The VTOL aircraft according to claim 6, wherein the electrically powered VTOL aircraft comprises a battery configured to supply power to the VTOL aircraft. [9] The VTOL aircraft according to claim 1, wherein: the first of the at least four variable-speed tilt rotors is mechanically coupled to the at least one wing on a first side of the fuselage; and the second of the at least four variable-speed tilt rotors is mechanically coupled to the at least one wing on a second side of the fuselage. [10] The VTOL aircraft according to claim 9, further comprising: a tail mechanically coupled to the fuselage, wherein: a third of the at least four variable-speed tilt rotors is mechanically coupled to the tail on the first side of the fuselage; and A fourth of at least four variable-speed tilting rotors is mechanically coupled to the tail on the second side of the fuselage. [11] The VTOL aircraft according to claim 10, wherein: the third of the at least four variable-speed tilt rotors is mechanically coupled to a first end of the tail on the first side of the fuselage; and The fourth of the at least four variable-speed tilting rotors is mechanically coupled to a second end of the tail on the second side of the fuselage. [12] The VTOL aircraft according to claim 11, wherein all of the at least four variable-speed tilting rotors are configured to tilt at least 70 degrees relative to the at least one wing. [13] The VTOL aircraft according to claim 11, further comprising: a first gondola mechanically coupled to the at least one wing on the first side of the fuselage and comprising a first static section fixed in relation to the at least one wing and a first tilting section tiltable in relation to the at least one wing, wherein the first of the at least four variable-speed tilting rotors is mechanically coupled to the first tilting section; a first battery designed to power the first motor; and a second battery designed to power the second motor, wherein the first battery is arranged in one of the first gondolas or at least one wing. [14] The VTOL aircraft according to claim 13, wherein the second battery is arranged in the other from the first gondola or the at least one wing. [15] The VTOL aircraft according to claim 13, wherein the first battery is arranged in the at least one wing. [16] The VTOL aircraft according to claim 15, wherein the first battery is longer in a direction perpendicular to the fuselage than in a direction parallel to the fuselage. [17] The VTOL aircraft according to claim 13, wherein the first battery is arranged in the first gondola. [18] The VTOL aircraft according to claim 17, wherein the first battery is longer in a direction parallel to the fuselage than in a direction perpendicular to the fuselage. [19] The VTOL aircraft according to claim 13, further comprising: a second gondola mechanically coupled to the at least one wing on the second side of the fuselage and comprising a second static section fixed in relation to the at least one wing and a second tilting section tiltable in relation to the at least one wing, wherein the second of the at least four variable-speed tilting rotors is mechanically coupled to the second tilting section; a third battery designed to power the third motor; and a fourth battery designed to power the fourth motor, wherein the third battery is arranged in one of the first gondolas or at least one wing. [20] The VTOL aircraft according to claim 19, wherein the fourth battery is arranged in the other from the second gondola or the at least one wing.

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