Fail-operational VTOL aircraft

Abstract

To provide a VTOL aircraft capable of vertical takeoff and landing even if one of multiple variable pitch rotor systems is inoperable.SOLUTION: In one aspect, described herein is an aircraft capable of carrying at least 181.437 kg of payloads. An embodiment has four rotors systems, each of the rotor systems being independently driven by an electric motor or other torque-producing source. Each of the rotor systems provides sufficient thrust such that the aircraft is capable of controlled vertical takeoff and landing even if one of variable pitch rotor systems is inoperable. An electronic control system is configured to control the rotational speed and pitch of at least one of the rotor systems.SELECTED DRAWING: Figure 9

Description

[Technical Field]

[0001] The field of this invention is vertical takeoff and landing (VTOL) aircraft. [Background technology]

[0002] VTOL aircraft have long been desired because they can take off and land with a small geometric footprint, thereby increasing the operational flexibility of aircraft, allowing vehicles to transport passengers and payloads closer to their desired destinations, and because they do not require the same infrastructure investment or land area as conventional takeoff and landing (CTOL) aircraft.

[0003] However, VTOL aircraft have several drawbacks compared to CTOL aircraft. Firstly, because VTOL aircraft require significantly more power to hover than CTOL aircraft require to take off via a runway, VTOL aircraft require significantly more on-board power (provided by the rated output of the on-board engines, drive motors, batteries, and drivetrain) than CTOL aircraft. The increased on-board power increases the weight and cost of the aircraft, limiting the vehicle's utility. Designs that are feasible with some powertrain configurations, such as battery-only power, may not be possible with certain VTOL configurations using current component technology due to limitations in battery and motor power density.

[0004] Secondly, VTOL aircraft typically produce more noise than CTOL aircraft of the same size, especially during takeoff and landing. This acoustic characteristic can be a public nuisance and may lead to operational restrictions on the use of vehicles, thus also reducing the usefulness of VTOL aircraft.

[0005] Finally, the safety record of VTOLs lags behind that of other parts of the aircraft industry. VTOL aircraft, and helicopters in particular, typically have one or more single points of failure in their rotor and powertrain systems. As a result, helicopters often recover from propulsion system failures by autorotation landings, which require a high level of pilot skill to control a good, but immediate landing. However, such landings become problematic in densely populated urban areas, where VTOL aircraft offer the greatest operational advantage in terms of landing site flexibility.

[0006] To address these shortcomings, a new type of VTOL aircraft, often called an electric-powered "eVTOL" aircraft, has emerged. Patent document 1 (Karem), referenced in this specification, examines the prior art. In the prior art, many vehicles use five or more rotors in VTOL flight, some of which can continue to operate safely even after the failure of a single rotor system. However, vehicles with more rotors inevitably have smaller diameter rotors with a smaller overall rotor disk area to fit within the same footprint, and therefore have a higher rotational load (weight divided by the total rotor disk area). This leads to increased noise during vertical takeoff and a decrease in the aircraft's horsepower load (thrust divided by the required power). A lower horsepower load reduces the vehicle's utility. A lower horsepower load increases the energy consumed by the aircraft and increases the required powertrain capacity. A large number of rotors would result in many lift rotors being exposed during forward flight, which would reduce the lift-to-drag ratio, thereby decreasing speed and range, or necessitate the use of numerous tilt rotors.

[0007] Karem addresses this horsepower load and noise issue by introducing vehicles with two to four rotor systems essential for flight. A four-rotor embodiment is shown in Figure 1. Such vehicles tend to have lower rotational loads within a given footprint (due to potentially larger rotor diameters), thereby gaining advantages in terms of horsepower load and noise, but these vehicles are less safe due to a single point of failure in the rotor system. Vehicles like Karem's attempt to improve safety with redundant motors, batteries, or engines, but still have a single point of failure in the reduction and rotor control systems.

[0008] Further prior art has attempted to address the hover horsepower load and noise drawbacks of VTOL aircraft with multiple coaxial rotors, and indeed, the noise and horsepower load advantages of coaxial rotors are known. As shown in Figure 2, Patent Document 2 discloses an aircraft comprising two rotor stacks, each having two rotor systems. The rotor systems of each stack are fixed relative to each other and cannot rotate independently, but are driven around a common axis, in order to increase the aircraft's level horsepower load and reduce noise. However, in each of the coaxial stacks, the rotors cannot rotate independently of each other.

[0009] As shown in Figure 3, Patent Document 3 discloses an aircraft having six stacks of two small coaxial propellers each for noise reduction and increased horsepower load. The two propellers in each coaxial stack can move relative to each other to adjust their azimuth to achieve optimal noise and horsepower load characteristics. While this design includes independent rotation for real-time adjustment of the azimuth spacing between propellers, it does not appear to be intended to oversize each rotor system to enable flight in the event that one of the two rotors in a rotor stack becomes inoperable. Furthermore, a person skilled in the art would not consider oversizing the rotors of the aircraft in Figure 3, because the aircraft has numerous backup propellers in case one fails, and the rotor diameters are relatively small and there are many rotors, so oversizing the thrust capacity of each rotor would not offer much benefit.

[0010] Figure 4 shows Airspace Experience Technology's MOBI-ONE, a tilt-wing aircraft featuring eight small rotors arranged in four coaxial stacks, each consisting of two rotors. The coaxial stacks rotate with the wings. The vehicle features a ninth rear auxiliary rotor for pitch control. The four coaxial stacks and the fifth auxiliary rotor allow for continued flight even if one coaxial stack becomes inoperable, but the increasing number of rotor stacks inevitably necessitates smaller rotors, which leads to higher rotational loads and consequently lower horsepower loads, as seen in some VTOL aircraft using small rotors.

[0011] Figures 5A and 5B show Embraer's PULSE concept, a tilt-rotor aircraft with four rotors arranged in two tilt-coaxial rotor stacks. Here, the coaxial design provides reduced noise and increased horsepower load compared to a single rotor with two tilt-rotors, but its relatively small rotors are sized to fit inside closed wings, which results in a lower horsepower load compared to aircraft with larger tilt-rotors. The PULSE concept discloses a rotor hover rotation speed of approximately 850 RPM, which follows conventional knowledge that rotors must rotate fast enough to avoid the gearbox. Conventional knowledge of urban air mobility also teaches that, in order to suppress noise, the rotor tip speed must be less than 137.16 meters per second (450 feet) when hovering, which would result in a rotor with a diameter of approximately 3.048 meters (10 feet). With this rotor size, the rotational load becomes high (146.47 kg / m 2 (30 lb / ft 2 (There is a high probability that this value is greater than or equal to), and at this value, it seems to the applicant that there is no prospect of a vehicle with a powertrain sized to provide all the thrust required by the vehicle by only one of the rotors in the coaxial stack. Therefore, while the PULSE concept aircraft may be able to land safely even if one rotor on each side becomes inoperable, it also seems that the aircraft would not be able to continue a normal flight plan, such as vertical takeoff.

[0012] Figure 6 shows a VTOL aircraft with four coaxial stacks, each consisting of two rotors, resulting in a total of eight rotors. In each coaxial stack, the rotors rotate in opposite directions to allow for torque control of the aircraft. However, this vehicle also uses numerous small rotors, in line with common knowledge. This aircraft may be able to descend smoothly by disconnecting power from one rotor to one of the coaxial stacks.

[0013] Figure 7 shows an eight-propeller aircraft as described in Patent Document 4 (Lyasoff). This application focuses on an embodiment having eight propellers. Figure 8 shows Joby's six-rotor VTOL. General knowledge would suggest that this design possesses a high level of reliability given by having a large number of rotors distributed around the aircraft. A person skilled in the art would not consider reducing the number of rotors in Joby's design, because this large number of rotors requires only a 150% oversizing of the propulsion system to cope with a situation where one rotor system fails and the opposing rotor system also stops to compensate for the balance of torque and moment. Modifying Joby's type of aircraft to have a significantly smaller number, but larger, rotors would also be contrary to general knowledge, as a person skilled in the art would consider that doing so would eliminate the additional reliability improvements that are presumably provided by the large number of rotors. Such a VTOL aircraft would also be contrary to the general knowledge of existing vehicles with two tilt-rotor or coaxial rotor stacks, which rely on the reliability of a single-point-of-failure rotor system that is sufficiently high to ensure the safety of the vehicle. [Prior art documents] [Patent Documents]

[0014] [Patent Document 1] U.S. Patent Application Publication No. 2018 / 334251 [Patent Document 2] U.S. Patent No. 8640985 [Patent Document 3] U.S. Patent Application Publication No. 2019 / 135413 [Patent Document 4] U.S. Patent Application Publication No. 2019 / 291862 [Overview of the Initiative] [Problems that the invention aims to solve]

[0015] Conventional technology presents a trade-off between configurations where VTOL flight can be sustained even with one rotor system inoperable (though with less desirable horsepower load, noise, and potentially complexity characteristics) and configurations where at least one rotor or drive system is essential for flight (though tend to have more desirable horsepower load, noise, and potentially complexity characteristics).

[0016] In particular with regard to coaxial VTOL designs, even when considering the use of a coaxial stack of propellers in conventional technology, the use of such rotors is largely consistent with conventional knowledge that it is easier to design and operate, often does not require a gearbox or similar mechanical complexity, and allows for redundancy by including extra rotors as needed, and uses smaller rotors.

[0017] Common knowledge teaches that when coaxial rotors are used, they result in increased horsepower load, reduced noise in the case of co-rotation, or improved torque balance in the case of counter-rotation. Sizing each rotor system in a coaxial stack (including associated reduction systems, engines, motors, and batteries) to provide all the thrust required for controlled flight in the absence of the other rotor (for example, in an aircraft with only two coaxial rotor stacks) contradicts common knowledge, as it would result in oversizing the rotor system by more than 200% to compensate for the lost thrust of the second rotor and overcome the additional inefficiency of a non-operating rotor in the rotor wake. Designers using coaxial rotor systems for increased horsepower load do not size each rotor system in a coaxial rotor set to be driven independently and to be able to provide all the thrust required for the set in the absence of one rotor, because this increases the weight and complexity of the system. A coaxial rotor set with a specific diameter and stiffness uses less total power to achieve a given thrust than a single rotor of the same diameter and stiffness would use for the same thrust. Furthermore, the power and torque required by each rotor system in a coaxial set is approximately half the total power of the rotor system. This negates the advantages that those skilled in the art perceive in using coaxial rotors to reduce the power required, as seen in some of the designs implemented in several embodiments herein. Oversizing a rotor system to enable VTOL flight in the event of a single rotor failure requires a much larger (not much smaller) settling power than a single rotor system.

[0018] What the prior art does not teach is a VTOL aircraft with less than five rotors stacked on two main coaxial rotor stacks or otherwise implemented, sized and dimensioned such that it has a low total rotational surface load and the aircraft can safely take off and land in a controlled state even when the rotor system is completely inoperable. Such an approach is contrary to the general knowledge that teaches that system redundancy usually requires a large number of five or more rotors to accommodate controlled VTOL flight after a complete failure of the rotor system, or that system redundancy results in an exorbitant weight. Instead, the vehicle design assumes that at least one rotor or drive system is essential for flight, thereby increasing the manufacturing and maintenance costs even assuming it is feasible, with a component-level reliability of about 10 -9 for failures per flight hour that require.

Means for Solving the Problem

[0019] In some embodiments presented herein, the VTOL aircraft has four main rotors arranged on two main coaxial rotor stacks or otherwise implemented, sized and dimensioned for the redundancy of each rotor system. In such a design, an essential rotor system may be replaced with a coaxial rotor set each having two or more independently driven rotor systems. In other embodiments, a non-coaxial rotor system in which the four main rotors are arranged in a quad configuration may also be utilized.

[0020] Unlike conventional coaxial rotors, which are primarily designed for anti-torque (without tail rotors), high edgewise advance ratios, or increased horsepower loads, some embodiments described herein provide coaxial rotor sets designed for redundancy and safety. Specifically, each coaxial rotor set is designed so that each individual rotor system within the set is driven independently and can provide all the thrust required for a coaxial rotor set that is essential for flight, including takeoff, in controlled VTOL flight.

[0021] In some coaxial embodiments, the rotor system is driven by a torque source (most likely a motor or engine), but this design may have a large excess of installed power if each torque source can only drive one of the rotors, or the system is designed so that each motor or engine can drive one rotor or the other depending on the mode of failure. The reduction system, and the motors or engines driving each rotor, are also sized to provide enough power for a single rotor in the stack to provide all the thrust required for the entire stack. If the aircraft is battery-powered, the battery system architecture provides sufficient battery power for the motors or other torque-generating devices driving the operational rotors. Furthermore, the aircraft has a way of balancing the aircraft's yaw moment when one of the rotor systems in the coaxial set is inoperable. In some embodiments, this method may involve stopping one or more rotor systems of other coaxial rotor sets, changing the speed and collective of the other coaxial rotor sets to change their respective yaw moment contributions, applying different rotor nacelle tilt angles to generate offset yaw moments, or other methods.

[0022] Some of the inventive concepts disclosed herein achieve improved levels of safety required for the air transport of people and cargo without the usual trade-offs of high rotational loads or increased rotor number, and this approach envisions large-diameter rotors without making those rotors and associated drivetrains unsafely essential.

[0023] Certain embodiments of the inventions disclosed herein may provide apparatus, systems, and methods for an electric vertical take-off and landing (eVTOL) aircraft designed to carry at least 400 pounds (approximately 181.437 kg) using two coaxial rotor stacks, wherein each stack consists of two variable-pitch rotor systems axially aligned around a substantially common axis of rotation, with each variable-pitch rotor preferably driven independently by one or more electric motors, but optionally by any other torque-generating device such as a fuel-consuming engine. In one embodiment, each variable-pitch rotor system is preferably a variable-speed stiff (non-articulated) rotor. Furthermore, these rotors are preferably tilt rotors configured to remain axially aligned with respect to each other while tilting with respect to the body. One or more of these rotors can be tilted to provide a significant amount of lift during rotor flight (e.g., vertical take-off) and forward thrust (or air braking) during wing flight. For clarity, in some embodiments, the relative azimuth angle between the two variable-pitch rotor systems in the coaxial stack is not fixed, and each variable-pitch rotor system can rotate independently of the other variable-pitch rotor system around a substantially common axis of rotation. Both variable-pitch rotor systems are configured to tilt together to the extent that the rotors tilt. The variable-pitch rotor may have a first blade, lbs / in at 30% of the radius at the root of the rotor. 2The flap stiffness of the first blade in a variable-speed variable-pitch rotor, expressed in units of feet, is at least 200 to the 4th power of the rotor diameter, expressed in feet, as in the rotor taught in U.S. Patent No. 6,641,365 (Karem), which is incorporated herein by reference. In some embodiments, the variable-pitch rotor is configured to operate at RPM levels less than 80% and even less than 60% of the maximum rotor system RPM level.

[0024] Some embodiments of the inventions disclosed herein are intended, in particular, for an aircraft designed to carry at least 181.437 kg (400 lbs), which is the payload capacity required to carry approximately two or more people or an equivalent amount of other payload. VTOL aircraft are subject to the square-cube scaling law, where an aircraft scaled to a larger size by a certain factor increases in rotor area, wing area, structural spar depth (which provides strength and stiffness), and other similar parameters by approximately the square of that factor, while weight increases by approximately the cube of that factor. The consequences have three aspects. First, the required structural strength, stiffness, and aeroelastic margins become more difficult to maintain as the scale increases. Second, the power required for hovering increases hyperlinearly as the overall size of the vehicle increases, although the aircraft's weight already increases hyperlinearly even with a given rotor-disk area. This means that the power margin required for propulsion redundancy increases undesirably with size. As a result, design techniques used in toys and smaller aircraft, particularly those designed to carry the equivalent of fewer than two people, tend to fail, especially in larger aircraft where the power requirements and design complexities are much greater. The inventive concepts contained herein may be particularly well suited to addressing the need for aircraft capable of carrying more than 181.437 kg (400 lbs). [Brief explanation of the drawing]

[0025] [Figure 1] A diagram illustrating a conventional four-rotor embodiment of a vehicle equipped with two to four rotor systems essential for flight. [Figure 2] A diagram showing a conventional aircraft. [Figure 3] A diagram showing an aircraft equipped with six stacks of two small, co-rotating propellers. [Figure 4] This diagram shows Airspace Experience Technology's MOBI-ONE, a tilt-wing aircraft equipped with eight small rotors arranged in four coaxial stacks, each consisting of two rotors. [Figure 5A] A diagram illustrating Embraer's PULSE concept, a tilt-rotor aircraft featuring four rotors arranged in two tilt-coaxial rotor stacks. [Figure 5B] An alternative diagram showing Embraer's PULSE concept, a tilt-rotor aircraft featuring four rotors arranged in two tilt-coaxial rotor stacks. [Figure 6] A diagram showing a VTOL aircraft equipped with four coaxial stacks, each consisting of two rotors, resulting in a total of eight rotors. [Figure 7] A diagram showing a conventional aircraft. [Figure 8] A diagram showing Joby's 6-rotor VTOL aircraft. [Figure 9] A perspective view of a preferred embodiment of a hovering VTOL aircraft according to the inventive concept of this specification. [Figure 10] A perspective view of a preferred embodiment of Figure 9, shown in a transition state. [Figure 11] A perspective view of a preferred embodiment of Figure 9, shown in cruising position. [Figure 12] A diagram showing possible embodiments of a drive system 900 for a two-rotor coaxial stack. [Figure 13] A diagram showing possible alternative embodiments of the drive system 900 for a two-rotor coaxial stack. [Figure 14]A diagram showing another possible alternative embodiment of the drive system 900 for a two-rotor coaxial stack. [Figure 15] A diagram showing possible alternative embodiments of the drive system 900 for a two-rotor coaxial stack. [Figure 16] A diagram showing one possible embodiment of a reversing mechanism for a drive system. [Figure 17] A diagram showing an embodiment in which the first wing is configured to pivot relative to the body of the aircraft. [Figure 18] A diagram showing an embodiment of an aircraft comprising at least a first auxiliary rotor system in addition to first and second coaxial rotor stacks. [Figure 19] A diagram showing an embodiment of an aircraft configured with four variable-pitch rotors, the aircraft being shown in a transitional state. [Figure 20] A diagram showing an embodiment of an aircraft configured with four variable-pitch rotors, the aircraft being shown in a hovering state. [Figure 21] A diagram showing an embodiment of an aircraft configured with four variable-pitch rotors, the aircraft being shown in a cruising position. [Figure 22] Figure 19 is a perspective view showing the aircraft. [Figure 23] A top view showing an embodiment of an aircraft configured with four variable-pitch rotors. [Figure 24] A diagram showing possible reduction systems, nacelles, and motor configurations that may be used (already described in U.S. Patent Application Publication No. 2018 / 334251 (Karem), but not in the context of the aircraft claimed herein). [Figure 25] A top view showing an embodiment of an aircraft configured with four variable-pitch rotors. [Figure 26] A diagram illustrating possible implementations of electronic flight control systems and related systems. [Figure 27] A diagram showing another possible implementation of an electronic flight control system and related systems. [Modes for carrying out the invention]

[0026] Figure 9 is a perspective view of a preferred embodiment of a hovering VTOL aircraft. The aircraft has a body 1101, an inner wing 1102, a tilting nacelle 1106, a tilting outer wing 1107, a first coaxial rotor stack 1103, and a first tail surface 1108. The outer wing 1107 tilts together with the tilting nacelle 1106 to reduce the downward load force from the rotor in hovering mode. The first coaxial rotor stack 1103 comprises a first variable-pitch rotor system 1104 and a second variable-pitch rotor system 1105, each rotor system rotatable around a substantially common axis of rotation, i.e., the rotor's axis of rotation, with each rotor system being driven independently by at least one torque source. The first and second variable-pitch rotors may be configured to rotate together or in opposite directions.

[0027] This embodiment further comprises a second coaxial rotor stack 1103. The second coaxial rotor stack 1103 comprises a first variable-pitch rotor system 1104 and a second variable-pitch rotor system 1105, each of which is rotatable about a substantially common axis of rotation, i.e., the axis of rotation of the rotor, with each rotor system being driven independently by at least one torque source.

[0028] The first variable-pitch rotor system 1104 includes rotor blades 1109. The rotor blades are of a rigid, hingeless type, such as those described in U.S. Patent No. 6,641,365 (Karem). Each variable-pitch rotor system independently provides thrust and force moment controlled by rotating the blades around a feather axis, such as in the rotor system described in U.S. Patent No. 1,035,1235 (Karem), incorporated herein by reference. An electronic control system 1411 simultaneously controls both the rotational speed and rotor blade pitch of each variable-pitch rotor system, as taught in U.S. Patent No. 6,641,365, to trim the rotor force and moment according to an electronic flight control system, and to ensure that the rotor operates at an optimal speed that yields an optimal blade lift distribution, and therefore with optimal efficiency and the minimum required rotor torque applied by the torque generating device. The inner blade 1102 transmits load from the coaxial rotor stack 1103 and the outer blade 1107 to the main body 1101.

[0029] Each of the first and second variable-pitch rotor systems in a coaxial rotor stack is capable of providing sufficient thrust to enable the aircraft to perform controlled vertical takeoffs and landings even if up to one of the variable-pitch rotor systems in the coaxial rotor stack is inoperable. This requires that the total thrust required for each coaxial stack, including the inefficiency of the additional rotor due to the inoperable rotor in the wake, be provided by a single rotor. To calculate the total required thrust, the maximum gross weight of the aircraft is taken into account, and then computational fluid dynamics (CFD) methods are used to calculate the additional thrust required to overcome the downward loads on the airframe caused by the wings, fuselage, nacelles, tail, and the inoperable rotor that may affect the rotor wake. Furthermore, an additional thrust margin is added for controllability and maneuverability in the event of one rotor being inoperable, although in a preferred embodiment such a thrust margin may exceed 15%. Next, a CFD method is used to calculate the rotor input torque and power required to achieve the desired thrust, including a margin, in the case of one rotor failure, under the desired operating conditions, including air density and rate of rise. Furthermore, this required rotor torque is used to size the torque and power ratings of each element of the powertrain, including the reduction system, torque generating device, and power distribution system (including a battery in embodiments where the torque generating device is an electric motor), and the power and torque ratings of each powertrain element further account for efficiency losses in each element of the powertrain system.

[0030] Each of the first variable-pitch rotor system 1104 and the second variable-pitch rotor system 1105 within the coaxial rotor stack 1103 may be sized such that each of the first and second variable-pitch rotor systems can provide all the thrust required by the electronic flight control system to the entire coaxial rotor stack even if one rotor is inoperable. To achieve this, the electronic control system may require increased RPM and power for the operational variable-pitch rotors. The increased RPM and power for the required thrust margin are calculated by computational fluid dynamics (CFD) methods for a particular rotor configuration, including the presence of an inoperable rotor, when sizing the rotor using results validated by physical rig tests of the rotor, nacelle, and adjacent wing surface combinations. During flight, the required RPM, blade pitch, and power are required by the electronic flight control system based on measured feedback of the aircraft's linear and angular velocities and acceleration for all three axes. Torque generating devices and reduction systems (shown in Figures 11, 12, and 13) are similarly sized with a sufficient margin in case one rotor becomes inoperable. If the aircraft is battery-powered, the batteries and distribution systems are sized to provide sufficient power to the operational rotors in each scenario in which a rotor becomes inoperable. Such sizing is typically done with a sufficient vertical climb rate of more than 30.48 m (100 ft) per minute, as well as the maximum total takeoff weight at the desired highest barometric altitude for operation with additional margins for control power, crosswinds, and trim penalties as defined by the desired flight envelope.

[0031] The fuselage 1101 is designed to carry payloads such as passengers, luggage, or cargo, and houses various systems including a landing gear configured as described in detail in U.S. Patent No. 10351235 (Karem), which is incorporated herein by reference. In a preferred embodiment, the offset of the vehicle's center of gravity from the center of the coaxial rotor stack is controlled by the mast moment from a rigid variable pitch rotor system, which necessarily requires cyclic control of the rotors.

[0032] In one exemplary embodiment, the aircraft of FIG. 9 has a maximum gross takeoff weight of 2721.554 kg (6000 lb). This aircraft has one cockpit and four passenger seats, and a luggage compartment. The fuselage of the aircraft is 10.0584 m (33 ft) in length, 1.524 m (5.0 ft) in maximum width, and 1.6764 m (5.5 ft) in maximum height. The wingspan of the aircraft is 13.716 m (45 ft), and the lateral distance between the centers of the two coaxial rotor stacks is 9.144 m (30 ft). The variable pitch rotor systems of the left and right coaxial rotor stacks have a diameter of 6.4008 m 2 (693 ft 2 ) with a non - overlapping rotor disk area and have a weight of 42.4771 kg / m 2 (8.7 lb / ft 2 ) based on the non - completely overlapping area, and 20.9944 kg / m 2 (4.3 lb / ft 2) has a rotational surface load. Each variable pitch rotor system is driven by four motors via a single-speed reduction system, and in hovering during nominal operating conditions, the motor input speed is 10,000 RPM and the output speed of the variable pitch rotor system is 400 RPM. However, any number of electric motors, such as one, two, or three, may be configured to drive each variable pitch rotor system via the reduction system. Each motor preferably has a maximum continuous output rating of 90 kW and a maximum emergency output rating of 110 kW, although the vehicle will typically use less than half of the maximum continuous output rating in hovering during most nominal operating conditions. The vehicle's 16 motors are powered by eight batteries, each capable of outputting 120 kW.

[0033] If one variable-pitch rotor system becomes inoperable, the electronic control system increases the speed of the remaining operational rotors to 533 RPM and increases the blade pitch to maintain the rotor operation near the rotor's peak efficiency thrust coefficient while providing the same thrust as in the nominal operating state. Importantly, the motor oversizing in some embodiments described herein allows the vehicle to maintain a positive performance margin in both vertical climb due to ground effect (at least 30.48 m / min (100 ft / min) climb) and control power (e.g., at least a 25% control power margin for vehicles operated in urban environments) at a barometric altitude of at least 1524 m (5000 ft). In the event of rotor system failure, the vehicles in some embodiments described herein allow the pilot to maintain controlled flight and transition to winged flight which requires much less power, and also allow the pilot to take off vertically and climb by ground effect if required, for example, in scenarios where landing is aborted. The same applies if the battery pack fails while all rotors are operating at their nominal capacity, and it is also possible to oversize the battery to anticipate a combined failure of the battery and rotor drive.

[0034] Figures 10 and 11 are perspective views of preferred embodiments in transition and cruising states, respectively, corresponding to the aircraft shown in Figure 9. As the tilting nacelle 1106 tilts the coaxial rotor stack 1103 around the tilt axis 1113 from a hovering position with a vertical thrust vector 1111 (shown in Figure 9) to a cruising position with a horizontal thrust vector 1112 (shown in Figure 11), the outer wing 1107 tilts from the vertical to the horizontal position at a similar angle of attack as the inner wing 1102. The outer wing increases the overall wing aspect ratio of the vehicle with reduced cruising drag. In a preferred embodiment, the main wing has an overall wing aspect ratio of at least 6, preferably greater than 8.

[0035] To achieve a desirable level of safe and efficient flight during hovering, transition, and cruising, the aircraft contemplated herein may be designed in accordance with the aerodynamic design teachings of U.S. Patent No. 1,0351,235 (Karem), which is incorporated in whole hereby. In such embodiments, the aircraft is designed to have a low wing stall speed of 80 KIAS or less, preferably 60 KIAS or less, which provides a wide transition corridor, thereby allowing the vehicle to tilt its rotors from hovering mode to cruising mode at low airspeeds. Such aircraft may also be advantageous because they may use slot wing flaps, as taught in U.S. Provisional Patent Application No. 62 / 757680 (Karem), which is incorporated hereby, to reduce the stall speed of the vehicle while providing a wing area that is efficient for higher-speed cruising. In some embodiments, the vehicle is designed to be efficient at cruising speed with a lift-to-drag ratio of at least 9, preferably at least 13, at a cruising speed of 130 kT.

[0036] Figures 12 and 13 are perspective views showing two embodiments of a drive system 900 for a two-rotor coaxial rotating stack, integrating a forward rotor system 901 and a rear rotor system 902 that rotate around a common shaft 903. Figure 12 shows a dual common assembly, and Figure 13 shows a back-to-back motor-reduction system array. The rotors can co-rotate by their ability to keep the available rotor stationary while either the forward or rear rotor continues to rotate when it becomes unavailable. The forward rotor hub 904 and rear rotor hub 905 are driven by the forward motor 906 or the rear motor 907 via a reduction gear set 908. The hubs 904 and 905 rotate on bearings 909 that are coaxially mounted on a structural nacelle extension 910. The collective and periodic pitch of the blades 911, which have blade shanks 912 located within feather bearings 913, may be controlled by rotary actuators 914 located within hubs 904 and 905, and electronic control of both motor speed and blade pitch is performed by a control and driver module 915. The drive system is lubricated and cooled by oil contained in tank 916, and independent spinner assemblies 917 and 918 surround these hub systems.

[0037] In one embodiment, such as the embodiment shown in Figure 14, the drive system 900 can be configured such that the front motor 906 and the rear motor 907 are torque-connected by a connecting shaft 919 that enables power sharing between the front variable-pitch rotor system and the rear variable-pitch rotor system. In such an embodiment, the drive system 900 is configured such that the rear motor 907 and the front motor 906 can drive either the front rotor hub 904 or the rear rotor hub 905, or both the rotor hub 904 and the rotor hub 905. In such a configuration, for example, if the rear reduction system 908 jams, the rear hub may not be driven, while the front hub is driven by the front and rear motors. Similarly, the rear hub may not be driven, while the rear hub is driven by the front and rear motors. Although not visible in Figure 24, a clutch such as a directional clutch 920a or 920b can be configured such that a first variable-pitch rotor system is driven when the forward and rear motors are driven in a first direction, and a second variable-pitch rotor system is driven when the forward and rear motors are driven in a second direction. Such an embodiment is preferable because the motors can be sized so that four motors are used to drive a single operable rotor system, rather than each of two sets of motors driving its own, thus easing the installation power requirements for driving only one of the variable-pitch rotor systems.

[0038] As shown in Figure 15, some embodiments may implement motors torque-connected to both rotors in a coaxial stack. Motors 906 and 907 may be inserted between sections of the connecting shaft 919. Directional clutches 920a and 920b may be inserted between each end of the connecting shaft 919 and their respective reduction systems. For example, in a reverse rotation embodiment, motor 906 may be configured to drive the connecting shaft 919 in a first direction. The front directional clutch 920a may be driven in engagement with the front reduction system and therefore with the front rotor. Under normal operating conditions, the rear directional clutch 920b connected to the same front motor 906 is configured to be disengaged. If the front rotor fails, the front motor 906 may reverse its direction of rotation, thereby disengaging from the front directional clutch 920a and engaging with the rear directional clutch 920b. Thus, even if one rotor becomes inoperable, all motors may be configured to generate thrust. Similarly, the rear motor 907 may be configured to drive the rear rotor during nominal operation and to drive the front rotor if the rear rotor becomes inoperable. Directional clutches 920a and 920b may be configured to provide desired nominal and failure mode operating characteristics. Although only three motors are visible in Figure 15, it should be understood that there may be six motors (two additional rear motors 907 and one additional front motor 906). Other embodiments may have any number of front and rear motors, such as two or four. The motors are shown between the front and rear rotor hubs, but may be positioned anywhere. Other implementations may include an alternative method for disconnecting the connecting shaft from the inoperable rotor, allowing it to drive a still-operable rotor, including the shear shaft.

[0039] Co-rotation embodiments may include a reversing gear, such as the reversing gear 1501 shown in Figure 16. The reversing gear is configured to allow the rear motor 907 and the front motor 906 to drive either the front hub 904 or the rear rotor hub 905. The reversing gear 1501 may have a pinion gear 1502, a reversing gear 1503, and a ring gear 1504, which are driven by the connected motor. The electric motor may be configured to drive the pinion gear 1502 such that the ring gear 1504 drives each hub in the opposite direction to how it would move if the reversing gear were not inserted.

[0040] In one embodiment shown in Figure 17, the first wing 2202 is configured to pivot relative to the body 2201. In some embodiments, the first wing 2202 may constitute a left wing 2203 and a right wing 2204, each configured to pivot substantially independently around a widthwise axis.

[0041] In another embodiment shown in Figure 18, the aircraft comprises, in addition to the first and second coaxial rotor stacks, at least a first auxiliary rotor system 1115, such as an auxiliary rotor disclosed in U.S. Patent Application Publication 2018 / 334251 (Karem), incorporated herein by reference. The aircraft may also comprise a second auxiliary rotor system 1115. In such embodiments, each of the first and second variable-pitch rotor systems of the coaxial rotor set and each of the first and second auxiliary rotor systems may work together to provide sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if up to one of the variable-pitch rotor systems of the coaxial rotor stack or up to one of the auxiliary rotor systems is inoperable. In one embodiment, the first auxiliary may have a rotor with a diameter of 50% or less of the diameter of the first variable-pitch rotor system. However, in some embodiments, the auxiliary rotor may have a diameter greater than 50%. Furthermore, each auxiliary rotor system may also be equipped with a variable-pitch rotor.

[0042] Figure 27 shows a preferred embodiment of the control law for an electronic flight control system in the case of a coaxial configuration. In some embodiments, four variable-pitch rotor systems may be arranged in a quad configuration, where the rotors are not coaxially aligned, but similar to the coaxial configuration, the variable-pitch rotor systems are sized and harmonized to provide an excess thrust margin so that the aircraft can perform flight even if one rotor is inoperable. Prior art teaches that at least five rotors are typically required to maintain vertical takeoff and landing with one rotor inoperable, but an aircraft with four variable-pitch rotor systems, if designed according to the teachings herein, can perform controlled vertical takeoff and landing even if up to one variable-pitch rotor system is inoperable.

[0043] A four-rotor configuration may be preferable because it allows for VTOL operation even when not all four rotors are operational, but it also anticipates the use of larger rotors. A fail-operational four-rotor tilt-rotor system, capable of VTOL flight even without using all rotors, may offer unexpected advantages.

[0044] In one exemplary embodiment shown in Figure 20, the aircraft is configured to have four variable-pitch rotors, namely a first variable-pitch rotor 1404, a second variable-pitch rotor 1404, a third variable-pitch rotor 1404, and a fourth variable-pitch rotor 1404. The aircraft has a body 1401, an inner wing 1402, a tilting nacelle 1406, a tilting outer wing 1407, an electronic flight control system 1411, and a tail surface 1408. Each variable-pitch rotor is configured to be driven by at least one torque source. As shown in Figure 23, the aircraft is configured such that the center of gravity 1409 is located approximately at the intersection of the first diagonal 1412 and the second diagonal 1413, where the first and second diagonals connect the centers of the variable pitch rotor thrusts 1403 of the first pair of variable pitch rotors 1404 and the second pair of variable pitch rotors 1404, respectively. It should be understood that the center of gravity 1409 may be located within an envelope surrounding the intersection of the first diagonal 1412 and the second diagonal 1413. Those skilled in the art will understand that the center of gravity 1409 moves within a small envelope based on the loads of passengers and payload, etc. Thus, the rotor positions relative to the aircraft's center of gravity are set so that the aircraft can perform controlled vertical takeoffs and landings even if up to one of the variable pitch rotors 1404 becomes inoperable. The electronic control system 1411 may be configured to control the rotational speed and blade pitch of the variable-pitch rotor. The variable-pitch rotor on the left side of the body may rotate together. Furthermore, the variable-pitch rotor on the right side may also rotate together. The variable-pitch rotor on the right side may rotate in the opposite direction to the variable-pitch rotor on the left side. The aircraft may have a second wing 1402. Preferably, at least one of the torque sources described above comprises an electric motor, but alternatively, a fuel-consuming engine may be provided.

[0045] In one particularly preferred embodiment, the aircraft in Figure 20 may have a maximum gross takeoff weight of 2721.554 kg (6,000 pounds) and may be configured to carry at least 181.437 kg (400 pounds). The aircraft has one cockpit and four passenger seats and a baggage compartment. The aircraft body is 10.668 m (35 ft) long, 1.524 m (5.0 ft) wide, and 1.6764 m (5.5 ft) high. The aircraft's wingspan is 14.9352 m (49 ft), and the widthwise distance between the centers of the two forward-mounted variable-pitch rotors is 8.8392 m (29 ft). The variable-pitch rotors are 116.7791 m 2 (1,257ft 2 ) has a rotor disc area of ​​6.096 m (20 ft) in diameter and a load capacity of 23.2892 kg / m 2 (4.77 lb / ft 2 The vehicle has a rotational surface load of ). Each variable-pitch rotor is driven by four motors via a single-speed reduction system, and in hovering during nominal operating conditions, the motor input speed is 10,000 RPM, and the output speed of the variable-pitch rotor system is 360 RPM. However, any number of electric motors, such as one, two, or three, may be configured to drive each variable-pitch rotor via the reduction system. Each motor has a maximum continuous output rating of 70 kW and a maximum emergency output rating of 90 kW, but the vehicle will typically use less than 40% of the maximum continuous output rating during hovering in most nominal operating conditions. The vehicle's 16 motors are powered by eight batteries, each capable of outputting at least 100 kW.

[0046] In such embodiments, the variable-pitch rotors may be sized so that none of the rotors are essential for flight. Failures per flight hour are 10 -9 For flight in urban environments where a certain level of system reliability is desired, the embodiment shown in Figure 20 may enable the aforementioned safety level at the aircraft level without requiring similar subsystem or component reliability.

[0047] Figure 21 shows an alternative diagram of the same embodiment as shown in Figure 20. The center of gravity 1409 is shown. The wing section 1402 is also shown. Figure 22 shows an alternative diagram of the same embodiment as shown in Figure 20. This embodiment is shown in winged flight mode. A thrust vector of 1405 is shown.

[0048] In one embodiment shown in Figure 23, controlled vertical takeoff and landing is achieved by reducing the power to the variable pitch rotors that are nearly opposite each other across an envelope containing the center of gravity 1409 from the non-operational variable pitch rotor. For example, if the forward left variable pitch rotor 1404 becomes inoperable due to failure, the electronic flight control system 1411 may reduce or cut off the power to the rear right variable pitch rotor 1404. Alternatively, the mast moment of the remaining variable pitch rotors may be used to compensate for the imbalance created by the non-operational rotor. A combination of mast moment control and thrust control may also be used to achieve controlled vertical takeoff and landing with fewer than all four variable pitch rotors. The variable pitch rotors 1404 and connected torque sources are configured to provide sufficient thrust to achieve vertical takeoff and landing when one or more variable pitch rotors are inoperable.

[0049] The reduction system, nacelle, and motor configurations shown in Figure 24 may be used. The drivetrain and nacelle configuration taught in U.S. Patent No. 10351235 (Karem), incorporated herein by reference, may also be used. In Figure 24, three motors 1905 drive the ring gear 1912. Such a system provides good further redundancy at the motor level. In the configuration shown in Figure 24, each variable-pitch rotor is driven by three motors 1905, but other numbers of motors are also possible. For example, two motors or four motors per rotor are possible. However, other drivetrain configurations may also be used with this embodiment.

[0050] A particularly preferred embodiment includes an electronic flight control system 1411 configured to respond to a fault without requiring a discontinuous mode switching based on fault diagnosis. Such an embodiment eliminates the risk of misdiagnosis of the aircraft's condition. One embodiment achieves this objective by dividing the functions of the electronic flight control system 1411 so that subsystems within the electronic flight control system handle fault conditions by architectural design rather than real-time determination.

[0051] In one embodiment shown in Figures 25 and 26, the electronic flight control system 1411 may group the first and second variable-pitch rotors 1404 as a first variable-pitch rotor pair 1414 when the vehicle is in VTOL mode. The electronic flight control system 1411 may group the third and fourth variable-pitch rotors 1404 as a second variable-pitch rotor pair 1414. In one preferred embodiment, the two variable-pitch rotors constituting each variable-pitch rotor pair are approximately opposite each other with respect to the center of gravity 1409. It should be understood that the two variable-pitch rotors comprising a variable-pitch rotor pair may be opposite each other with respect to the envelope encompassing the center of gravity 1409. Similarly, the third and fourth variable-pitch rotors 1404 are opposite each other with respect to the envelope containing the center of gravity.

[0052] As shown in Figure 25, the horizontal plane is defined by the aircraft's roll axis 1422 and pitch axis 1416. The x-axis 1420 of the variable-pitch rotor pair runs through an envelope encompassing the center of gravity 1409, between the centers of the variable-pitch rotors comprising each variable-pitch rotor pair 1414. The coordinate frame of the first variable-pitch rotor pair has an x-axis 1420 intersecting with the center of the variable-pitch rotor comprising the first variable-pitch rotor pair. The coordinate system of the first variable-pitch rotor pair further comprises the y-axis 1415 of the corresponding variable-pitch rotor pair, which is orthogonal to the x-axis and runs parallel to the horizontal plane. The coordinate system of the first variable-pitch rotor further comprises the z-axis of the first variable-pitch rotor pair, which is orthogonal to both the x-axis and y-axis of the variable-pitch rotor and may also correspond to altitude. The coordinate system for the first variable-pitch rotor further comprises a first moment axis 1419 around the x-axis of the variable-pitch rotor pair, and a second moment axis 1423 around the y-axis 1415 of the variable-pitch rotor pair. The coordinate frame for the second variable-pitch rotor pair has an x-axis 1420 intersecting the center of the variable-pitch rotor comprising the second variable-pitch rotor pair 1414. The coordinate system for the second variable-pitch rotor pair further comprises a y-axis 1415 of the corresponding variable-pitch rotor pair, orthogonal to the x-axis 1420 and running parallel to the horizontal plane. The coordinate system for the second variable-pitch rotor further comprises a z-axis of the first variable-pitch rotor pair, orthogonal to both the x-axis and y-axis of the variable-pitch rotor, and may also correspond to it. The coordinate system for the second variable-pitch rotor further comprises a first moment axis 1419 around the x-axis of the variable-pitch rotor pair, and a second moment axis 1423 around the y-axis 1415 of the variable-pitch rotor pair.

[0053] The first variable-pitch rotor pair 1414 can generate a net thrust in the z-axis direction and a moment around the y-axis 1415 of each variable-pitch rotor pair. The vertical thrust of each variable-pitch rotor pair 1414 is the sum of the thrusts of each variable-pitch rotor 1404 in each variable-pitch rotor pair 1414. The moment around the y-axis of each variable-pitch rotor pair is obtained by multiplying the thrust difference between the two rotors comprising each variable-pitch rotor pair 1414 by the distance between the centers of each variable-pitch rotor.

[0054] As shown in Figure 26, the electronic flight control system 1411 may have two command channels for each variable pitch rotor pair 1414: a variable pitch rotor pair thrust signal 2305 for requesting the net thrust of the pair, and a variable pitch rotor pair moment signal 2304 for requesting the moment of the first variable pitch rotor pair. The signal values ​​for each of these channels are calculated by separate control laws. The variable pitch rotor pair thrust control law 2303 adjusts the vehicle altitude and generates commands for the variable pitch rotor pair thrust signal, while the attitude control law 2302 adjusts the attitude and generates commands for the variable pitch rotor pair moment signal 2304.

[0055] The electronic flight control system 1411 may receive vehicle dynamics values ​​from the vehicle dynamics sensor 1421. The vehicle dynamics sensor 1421 may comprise one or more of the following known sensors: GPS, magnetometer, IMU, or other. The electronic flight control system may use the values ​​from the vehicle dynamics sensor 1421 to calculate a vehicle dynamics error value. The vehicle dynamics error value is converted into parameters corresponding to the coordinate frames of the first and second variable-pitch rotor pairs. The electronic flight control system 1411 may proportionally increase the vehicle dynamics error value for a given parameter. A system attitude input is generated. This system attitude input is added to or subtracted from the corresponding system altitude input. The combined input is then transmitted to the vehicle dynamics control device. In Figure 26, these signals are shown as being transmitted to vehicle dynamics control devices associated with each rotor, such as rotor blade pitch actuators and motors; however, it should be understood that these signals may also be transmitted to other vehicle dynamics control devices. Vehicle dynamics control devices may include an electric motor coupled to one of the variable-pitch rotors, an actuator configured to control the blade pitch of a variable-pitch rotor, a control surface actuator, a rotor tilt actuator, or any other device configured to control vehicle dynamics.

[0056] In such preferred embodiments, such as the embodiment shown in Figure 25, failure of one variable-pitch rotor 1404 in a variable-pitch rotor pair 1414 impairs the ability of each variable-pitch rotor pair to independently generate net thrust and moment. However, as long as the other variable-pitch rotor in the variable-pitch rotor pair 1414 is functioning, that pair retains its ability to generate a combination of thrust and moment. In the event of a failure in which one rotor becomes inoperable, the failed variable-pitch rotor pair continues to adjust the vehicle attitude but cannot contribute to adjusting the vehicle altitude. The other variable-pitch rotor pair 1414, consisting of two functioning variable-pitch rotors, then becomes responsible for adjusting the altitude.

[0057] The electronic flight control system 1411 may include a reference table of thrust coefficients for the collective. The electronic flight control system may use the reference table to determine preferred RPM and collective pitch settings for a perceived flight condition. Configuring the electronic flight control system 1411 with the reference table allows for the use of ideal RPM and collective pitch for a given flight condition without the input of additional measurements, thereby enabling the aircraft to exhibit ideal flight characteristics, such as low noise and high efficiency.

[0058] The control system maintains the vehicle's altitude and attitude. One advantage of such an embodiment is that the electronic flight control system can be configured to provide continuous and smooth control in the event of a component failure in the aircraft.

[0059] If one rotor becomes inoperable, the electronic flight control system 1411 may reduce or cut off power to one of the operational variable-pitch rotors. The electronic flight control system 1411 may also change the mast moment of the operational variable-pitch rotors to bring about a force merge that allows the aircraft to perform vertical takeoff and landing even when not all of the variable-pitch rotors 1404 are operational. Alternatively, the electronic flight control system 1411 may reduce power to at least one variable-pitch rotor and change the pitch moment of the operational variable-pitch rotors 1404 to bring about a desired force merge.

[0060] Similarly, an aircraft of one embodiment contemplated herein may be configured to achieve VTOL flight with two of its variable-pitch rotors inoperable when a non-operable variable-pitch rotor is located at a diagonal corner. For example, in Figure 23, if both the front left variable-pitch rotor 1404 and the rear right variable-pitch rotor 1404 become inoperable, VTOL flight may be achieved using only the front right variable-pitch rotor 1404 and the rear left variable-pitch rotor 1404.

[0061] It should be noted that any expression relating to an electronic flight control system, or control and driver modules, should be read as including any appropriate combination of computing devices, such as servers, interfaces, systems, databases, agents, peers, engines, control devices, or other types of computing devices operating individually or collectively. These computing devices may also include processors configured to execute software instructions stored in tangible, non-temporary computer-readable storage media (e.g., hard drives, solid-state drives, RAM, flash memory, ROM, etc.). The software instructions are preferably configured to provide the computing devices with respect to the disclosed devices the roles, responsibilities, or other functions described above. In some embodiments, various servers, systems, databases, or interfaces may exchange data using standardized protocols or algorithms, possibly based on HTTP, HTTPS, AES, public / private key exchange, web service APIs, known financial transaction protocols, or other electronic information exchange methods. Data exchange is preferably carried out over a packet-switched network, the Internet, a LAN, a WAN, a VPN, or other type of packet-switched network. The electronic flight control system may be located anywhere on the aircraft where the actuator is located, or it may be located anywhere else, such as in a ground control center, on another aircraft, or even within a component of the actuator itself.

[0062] The electronic flight control system may be located anywhere on the aircraft, or at any other location, such as in a ground control center or on another aircraft. Furthermore, in some embodiments, the electronic flight control system, as well as the control and driver modules, may be implemented in separate units or combined into a single unit.

[0063] It should be understood that the concepts taught herein are equally applicable to propellers, rotors, or prop-rotors, whether used in helicopters, airplanes, or tilt-rotor aircraft. The term rotor shall be understood to encompass rotors, propellers, and prop-rotors, or any other rotating wing configured to generate thrust and / or lift. Furthermore, rotor blades shall be understood to encompass any blades belonging to rotor blades, propeller blades, and prop-rotor blades, or any other rotating wing configured to generate thrust and / or lift. For example, some embodiments can be used to implement independent blade control for turboprops. Such implementations can reduce noise and vibration in various flight conditions.

[0064] While several embodiments are described herein, they should be understood to be illustrative and not limiting in any way. Furthermore, it should be understood that aspects of various embodiments may be combined with aspects of other embodiments. The technical concepts included in this disclosure are described below. (Note 1) An aircraft capable of carrying a payload of at least 181.437 kg (400 pounds), The main unit and The system comprises a first variable-pitch rotor, a second variable-pitch rotor, a third variable-pitch rotor, and a fourth variable-pitch rotor, each of which is independently driven by at least one torque source. Each of the first, second, third, and fourth variable-pitch rotors is a variable-speed rotor, An aircraft in which each of the first variable-pitch rotor, the second variable-pitch rotor, the third variable-pitch rotor, and the fourth variable-pitch rotor provides sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even when only three of the variable-pitch rotors are operational. (Note 2) The aircraft according to Appendix 1, further comprising a first wing mechanically coupled to the main body. (Note 3) The aircraft as described in Appendix 2, further comprising a second wing mechanically coupled to the aforementioned body. (Note 4) The aircraft according to Appendix 2, wherein the first wing is configured to pivot relative to the body. (Note 5) The aircraft according to Appendix 1, wherein the first and second variable-pitch rotors are located on opposite sides of an envelope containing the center of gravity, and the first and second variable-pitch rotors can work together to provide all the thrust necessary to enable the aircraft to perform controlled vertical takeoffs and landings. (Note 6) The aircraft according to Appendix 1, further comprising a tilt mechanism configured to tilt the first variable-pitch rotor by at least 80 degrees. (Note 7) The aircraft according to Appendix 1, wherein the first variable-pitch rotor and the second variable-pitch rotor are located on the port side of the body, the third variable-pitch rotor and the fourth variable-pitch rotor are located on the starboard side of the body, and the first and second variable-pitch rotors are configured to rotate in a first direction. (Note 8) The aircraft according to Appendix 7, wherein the third and fourth variable-pitch rotors are configured to rotate in a second direction, the second direction being opposite to the first direction. (Note 9) The aircraft according to Appendix 1, wherein the first variable-pitch rotor is configured to provide rotor circulation control. (Note 10) The aircraft according to Appendix 9, wherein the first variable-pitch rotor comprises rigid blades and a hub configured to apply a mast moment to the aircraft. (Note 11) The aircraft according to Appendix 10, wherein the first variable-pitch rotor comprises rotor blades that are individually controlled. (Note 12) The aircraft described in Appendix 1, wherein the aircraft is equipped with four or fewer variable-pitch rotors. (Note 13) The aircraft as described in Appendix 1, wherein the first torque source is a first electric motor. (Note 14) The aircraft according to Appendix 13, further comprising a reduction system for operably coupling the first variable-pitch rotor and the first electric motor. (Note 15) The aircraft according to Appendix 14, wherein the first variable-pitch rotor is driven by the first electric motor and at least the second electric motor via the reduction system. (Note 16) The aircraft described in Appendix 1, wherein the first variable-pitch rotor is configured to operate at an RPM level less than 60% of the maximum rotor RPM level. (Note 17) The aircraft according to Appendix 1, wherein the electronic flight control system is configured to use vehicle attitude data to request controlled flight of the aircraft even when only three of the rotors are operational. (Note 18) The aircraft according to Appendix 1, wherein the electronic flight control system is configured to use vehicle attitude data to request controlled flight of the aircraft even when only three of the rotors are operational, without using discontinuous state switching. (Note 19) A method for operating a VTOL aircraft, The process of receiving commands regarding the vehicle's attitude, A process of requesting thrust for one of the four variable-pitch rotors, The process of receiving vehicle dynamics signals from a vehicle dynamics sensor, The process of determining, using the received vehicle dynamics signal, that the thrust required for the variable-pitch rotor is insufficient to achieve the required vehicle attitude, The process involves increasing the thrust required for each of the variable-pitch rotors until sufficient thrust is generated to perform controlled vertical takeoff and landing, A method that includes [a certain feature]. (Note 20) The method according to Appendix 19, further comprising the additional step of generating sufficient thrust for the aircraft to perform controlled vertical takeoff and landing even when only three of the variable-pitch rotors are operational. (Note 21) An aircraft capable of carrying a payload of at least 181.437 kg (400 pounds), The main unit and A first and second coaxial rotor stack, each having a first variable-pitch rotor system and a second variable-pitch rotor system rotatable about a substantially common axis of rotation, each of which rotor systems is independently driven by at least a first torque source, The system comprises an electronic control system configured to control the rotational speed and blade pitch of at least the first variable-pitch rotor system of each of the first and second coaxial rotor stacks, An aircraft wherein each of the first and second variable-pitch rotor systems of the coaxial rotor stack provides sufficient thrust to enable the aircraft to perform controlled vertical takeoffs and landings even if up to one of the variable-pitch rotor systems of the coaxial rotor stack is inoperable. (Note 22) The aircraft described in Appendix 21, further comprising wings mechanically coupled to the aforementioned body. (Note 23) The aircraft according to Appendix 22, wherein the wings are configured to pivot relative to the main body. (Note 24) The aircraft according to Appendix 23, wherein the wing is divided into a left wing and a right wing, and each of the left wing and the right wing is configured to pivot independently about a substantially widthwise axis. (Note 25) The aircraft according to Appendix 21, further comprising a tilt mechanism configured to tilt the first coaxial rotor stack by at least 80 degrees. (Note 26) The aircraft according to Appendix 21, wherein the first and second variable-pitch rotor systems of the first coaxial rotor stack are configured to rotate in the same direction. (Note 27) The aircraft according to Appendix 21, wherein the first and second variable-pitch rotor systems of the first coaxial rotor stack are configured to rotate in opposite directions. (Note 28) The aircraft according to Appendix 21, wherein the first variable-pitch rotor system of the first coaxial rotor stack is configured to provide rotor circulation control. (Note 29) The aircraft according to Appendix 28, wherein the first variable-pitch rotor system of the first coaxial rotor stack comprises rigid blades and hubs configured to apply a mast moment to the aircraft. (Note 30) The aircraft according to Appendix 28, wherein the first variable-pitch rotor system of the first coaxial rotor stack comprises rotor blades that are individually controlled. (Note 31) The aircraft described in Appendix 21, wherein the first torque source is a fuel-consuming engine. (Note 32) The aircraft as described in Appendix 31, wherein the first torque source is a first electric motor. (Note 33) The aircraft according to Appendix 32, further comprising a reduction system for operably coupling the first variable-pitch rotor system and the first electric motor. (Note 34) The aircraft according to Appendix 33, wherein the first variable-pitch rotor system of the first coaxial rotor stack is driven by the first electric motor and at least the second electric motor via the reduction system. (Note 35) The aircraft according to Appendix 31, further comprising a first mechanical clutch system, wherein at least the first torque source configured to drive the first variable-pitch rotor system of the first coaxial rotor stack when the first mechanical clutch system is disengaged is configured to drive the second variable-pitch rotor system of the first coaxial rotor stack. (Note 36) The aircraft as described in Appendix 31, wherein the first and second variable-pitch rotor systems of the first and second coaxial rotor stacks, together with the torque sources and the electronic control system, are sized and configured to enable controlled vertical takeoff and landing without the use of additional lift sources, even if up to one of the variable-pitch rotor systems of the coaxial rotor stacks becomes inoperable. (Note 37) The aircraft as described in Appendix 31, wherein the first variable-pitch rotor system is configured to operate at an RPM level less than 80% of the maximum rotor system RPM level. (Note 38) The aircraft described in Appendix 37, wherein either the first or second variable-pitch rotor is configured to operate at an RPM level less than 60% of the maximum rotor system RPM level. (Note 39) An aircraft capable of carrying a payload of at least 181.437 kg (400 pounds), The main unit and A first and second coaxial rotor stack, each having a first variable-pitch rotor system and a second variable-pitch rotor system rotatable around a substantially common axis of rotation, each of which rotor systems is independently driven by at least one torque source, An electronic control system configured to control the rotational speed and pitch of at least the first variable-pitch rotor system of each of the first and second coaxial rotor stacks, A first auxiliary rotor system not housed in the first or second coaxial rotor stack, comprising: An aircraft in which each of the first and second variable-pitch rotor systems of the coaxial rotor set and the first auxiliary rotor system cooperate to provide sufficient thrust to enable the aircraft to perform controlled vertical takeoffs and landings even if up to one of the variable-pitch rotor systems or up to one of the auxiliary rotor systems of the coaxial rotor stack is inoperable. (Note 40) The aircraft described in Appendix 39, further comprising a second auxiliary rotor.

Claims

1. An aircraft capable of carrying a payload of at least 181.437 kg (400 pounds), The main unit and Mechanically coupled to the main body, at least one wing having an axis in the width direction with respect to the main body, A first and second coaxial rotor stack, each having first and second variable-pitch rotor systems rotatable around a substantially common axis of rotation, each of which rotor systems is independently driven by at least a first torque source, The system comprises an electronic control system configured to control the rotational speed and blade pitch of at least the first variable-pitch rotor system of each of the first and second coaxial rotor stacks, Each of the first and second coaxial rotor stacks is configured to be inclined with respect to the width axis, An aircraft wherein each of the first and second variable-pitch rotor systems of the coaxial rotor stack provides sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if up to one of the variable-pitch rotor systems of the coaxial rotor stack is inoperable.

2. The aircraft according to claim 1, wherein at least one of the first and second coaxial rotor stacks is mechanically coupled to the at least one wing.

3. The aircraft according to claim 2, wherein the at least one wing is configured to pivot relative to the body.

4. The aircraft according to claim 3, wherein the at least one wing comprises a left wing and a right wing, and each of the left wing and the right wing is configured to pivot independently about the width axis.

5. The aircraft according to claim 1, further comprising a tilt mechanism configured to tilt the first coaxial rotor stack by at least 80 degrees.

6. The aircraft according to claim 1, wherein the first and second variable-pitch rotor systems of the first coaxial rotor stack are configured to rotate in the same direction.

7. The aircraft according to claim 1, wherein the first and second variable-pitch rotor systems of the first coaxial rotor stack are configured to rotate in opposite directions.

8. The aircraft according to claim 1, wherein the first variable-pitch rotor system of the first coaxial rotor stack is configured to provide rotor circulation control.

9. The aircraft according to claim 8, wherein the first variable-pitch rotor system of the first coaxial rotor stack comprises rigid blades and hubs configured to apply a mast moment to the aircraft.

10. The aircraft according to claim 8, wherein the first variable-pitch rotor system of the first coaxial rotor stack comprises rotor blades that are individually controlled.

11. The aircraft according to claim 1, wherein the first torque source is a fuel-consuming engine.

12. The aircraft according to claim 1, wherein the first torque source is a first electric motor.

13. The aircraft according to claim 12, further comprising a reduction system for operably coupling the first variable-pitch rotor system and the first electric motor.

14. The aircraft according to claim 13, wherein the first variable-pitch rotor system of the first coaxial rotor stack is driven by the first electric motor and at least the second electric motor via the reduction system.

15. The aircraft according to claim 1, further comprising a first mechanical clutch system, wherein at least one first torque source configured to drive the first variable-pitch rotor system of the first coaxial rotor stack when the first mechanical clutch system is disengaged is configured to drive the second variable-pitch rotor system of the first coaxial rotor stack.

16. The aircraft according to claim 1, wherein the first and second variable-pitch rotor systems of the first and second coaxial rotor stacks, together with at least the first torque source and the electronic control system, are sized and configured to enable controlled vertical takeoff and landing without the use of additional lift sources, even if up to one of the variable-pitch rotor systems of the coaxial rotor stacks becomes inoperable.

17. The aircraft according to claim 1, wherein the first variable-pitch rotor system is configured to operate at an RPM level less than 80% of the maximum rotor system RPM level.

18. The aircraft according to claim 17, wherein either the first or second variable pitch rotor system is configured to operate at an RPM level less than 60% of the maximum rotor system RPM level.

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