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 the invention is vertical takeoff and landing (VTOL) aircraft. [Background technology]

[0002] VTOL aircraft have long been desirable because they can take off and land in a small geometric footprint, thereby increasing aircraft operational flexibility by allowing vehicles to carry passengers and payload closer to desired destinations, and without requiring the same infrastructure investment or land area required by conventional takeoff and landing (CTOL) aircraft.

[0003] However, VTOL aircraft have several disadvantages compared to CTOL aircraft. First, because a VTOL aircraft requires significantly more power to hover than a CTOL aircraft to take off via a runway, a VTOL aircraft requires significantly more installed power (given by the installed engines, drive motors, batteries, and drivetrain power ratings) than a CTOL aircraft. The increased installed power increases the weight and cost of the aircraft, limiting the usefulness of the vehicle. Designs that are feasible for some powertrain configurations, such as battery-only power, may not be possible for certain VTOL configurations using current component technology due to limitations in the power density of the batteries and motors.

[0004] Second, VTOL aircraft are typically louder than comparably sized CTOL aircraft, especially during takeoff and landing. This acoustic feature again reduces the usefulness of VTOL aircraft, as it can be a public nuisance and lead to operational restrictions on the use of the vehicles.

[0005] Finally, VTOL's safety record lags behind the rest of the aviation industry. VTOL aircraft, and helicopters in particular, typically have one or more single points of failure in their rotors 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, but such landings are problematic in densely populated urban areas where VTOL aircraft offer the greatest operational advantage in terms of landing location flexibility.

[0006] To address these shortcomings, a new class of VTOL aircraft, often referred to as electric "eVTOL" aircraft, is emerging. US Patent No. 5,999,943 (Karem), incorporated herein by reference, surveys the prior art. In the prior art, many vehicles use five or more rotors in VTOL flight, some of which can continue safe operation after failure of a single rotor system. However, vehicles with more rotors will necessarily have smaller diameter rotors with a smaller total vehicle rotor disk area to fit into the same footprint, and therefore higher rotating surface loading (weight divided by total rotor disk area). This leads to increased noise during vertical takeoff and a lower horsepower loading (thrust divided by required power) of the aircraft. A lower horsepower loading reduces the usefulness of the vehicle. A lower horsepower loading increases the energy consumed by the aircraft and increases the powertrain capacity requirements. Multiple rotors would also result in more lifting rotors being exposed in forward flight, which would reduce the lift to drag ratio and therefore reduce speed and range, or would require multiple tilting rotors.

[0007] Karem addresses this horsepower load and noise issue by introducing vehicles with two to four flight-critical rotor systems. A four-rotor embodiment is shown in Figure 1. Such vehicles tend to have lower rotating surface loads in a given footprint (due to potentially larger rotor diameters), thereby garnering advantages in terms of horsepower load and noise, but these vehicles are less safe due to single points of failure in the rotor system. Vehicles such as Karem's attempt to increase safety with redundant motors, batteries, or engines, but still have single points of failure in the retarding and rotor control systems.

[0008] Further prior art attempts to address the hover power load and noise drawbacks of VTOL aircraft with multiple co-rotating coaxial rotors, and indeed the noise and power load advantages of coaxial rotors are known. As shown in FIG. 2, US Pat. No. 5,399,366 discloses an aircraft with two rotor stacks, each with two rotor systems. The rotor systems of each stack are fixed relative to each other, cannot rotate independently, and are driven about a common axis for increased level power load of the aircraft and reduced noise. However, in each of the coaxial stacks, the rotors cannot rotate independently of each other.

[0009] As shown in FIG. 3, the patent application 2003 / 0133633 discloses an aircraft with six stacks of two small co-rotating propellers for reduced noise and increased horsepower load. The two propellers of each coaxial stack can move relative to each other to adjust the azimuth angle for optimal noise and horsepower load characteristics. Although the design includes independent rotation for real-time adjustment of the azimuth spacing between the propellers, it does not appear to contemplate oversizing each rotor system to allow flight with one of the two rotors of the rotor stack inoperable. Furthermore, a person skilled in the art would not consider oversizing the rotors of the aircraft of FIG. 3. The relatively small rotor diameter and large number of rotors make oversizing the thrust capacity of each rotor less advantageous since the aircraft has many backup propellers in case one fails.

[0010] Figure 4 shows Airspace Experience Technology's MOBI-ONE, a tilt-wing aircraft with eight small rotors arranged in four co-rotating coaxial stacks of two rotors each. The coaxial stacks rotate with the wing. The vehicle features a ninth aft auxiliary rotor for pitch control. Although a total of four coaxial stacks and a fifth auxiliary rotor allow for continued flight in the event one coaxial stack is inoperable, the increased number of rotor stacks necessarily forces the rotors to be smaller, which leads to the higher rolling surface loads and resulting lower horsepower loads found in VTOL aircraft using some small rotors.

[0011] 5A and 5B show Embraer's PULSE concept, a tilt rotor aircraft with four rotors arranged in two tilting coaxial rotor stacks. Here, the coaxial design offers reduced noise and increased horsepower loading over two tilting single rotors, but the relatively small rotors are sized to fit inside closed wings, which results in lower horsepower loading compared to aircraft with larger tilting rotors. The PULSE concept discloses a rotor hover rotation speed of about 850 RPM, which follows conventional wisdom that rotors must rotate fast enough to avoid gearboxes. Conventional wisdom in urban air mobility also teaches that to suppress noise, rotor tip speed must be less than 450 feet per second in hover, which would result in a rotor with a diameter of about 10 feet. This rotor size has a high surface load (146.47kg / m 2 (30lb / ft 2 ) or higher), which appears to the applicant to make it unlikely that a vehicle with a powertrain sized to provide all the thrust required by the vehicle through only one of the rotors in the coaxial stack would be possible. Thus, while a PULSE concept aircraft might be able to land safely if one rotor on each side were to become inoperable, it also appears that the aircraft would not be able to continue with its normal flight plan, e.g., vertical takeoff.

[0012] Figure 6 shows a VTOL aircraft with four coaxial stacks of two rotors each for a total of eight rotors. In each coaxial stack, the rotors rotate in opposite directions, allowing torque control of the aircraft. However, this vehicle also uses multiple smaller rotors, in keeping with conventional wisdom. It is possible that the aircraft could successfully descend with power cut to one rotor from each coaxial stack.

[0013] Figure 7 shows an eight-propeller aircraft described in US Patent No. 5,399,633 (Lyasoff), which focuses on an embodiment with eight propellers. FIG. 8 shows a Joby six-rotor VTOL. Conventional wisdom would suggest that this design has a high degree of reliability imparted by having multiple rotors distributed around the aircraft. One skilled in the art would not consider reducing the number of Joby rotors because the multiple rotors would only require a 150% oversizing of the propulsion system to accommodate a situation where a rotor system fails and the opposing rotor system is also shut down to compensate for torque and moment balance. Modifying a Joby type aircraft to have a significantly fewer, but larger, rotors would also go against conventional wisdom, as one skilled in the art would consider doing so would eliminate the additional reliability improvement presumably imparted by the multiple rotors. Such a VTOL aircraft would also go against conventional wisdom of existing vehicles with two tilting rotors or coaxial rotor stacks, which rely on the vehicle having a high enough reliability in the rotor system with no single point of failure to ensure safety. [Prior art documents] [Patent documents]

[0014] [Patent Document 1] US Patent Application Publication No. 2018 / 334251 [Patent Document 2] U.S. Patent No. 8,640,985 [Patent Document 3] US Patent Application Publication No. 2019 / 135413 [Patent Document 4] US Patent Application Publication No. 2019 / 291862 Summary of the Invention [Problem to be solved by the invention]

[0015] In the prior art, there is a trade-off between a configuration in which VTOL flight can be sustained with one rotor system inoperable (but having less desirable power load, noise, and sometimes complexity characteristics) and a configuration in which at least one rotor or drive system is essential to flight (but which tends to have more desirable power load, noise, and sometimes complexity characteristics).

[0016] With respect to coaxial VTOL designs specifically, even where the prior art contemplates the use of a coaxial stack of propellers, the use of such rotors is largely consistent with conventional wisdom of using smaller rotors which are easier to design and operate, often do not require gearboxes or similar mechanical complexities, and allow for redundancy by including extra rotors as needed.

[0017] Conventional wisdom teaches that when coaxial rotors are used, they provide increased horsepower loading and reduced noise in the case of co-rotation or improved torque balance in the case of counter-rotation. Sizing each rotor system (including associated reduction systems, engines, motors, and batteries) in a coaxial stack to provide all of the thrust required for controlled flight in the absence of the other rotor (e.g., in an aircraft with only two coaxial rotor stacks) is contrary to conventional wisdom because it would result in oversizing the rotor systems by more than 200% to compensate for the lost thrust of the second rotor and overcome the additional inefficiency of an inoperative rotor in the rotor wake. Designers who use coaxial rotor systems for increased horsepower loading do not size each rotor system of a coaxial rotor set to be independently driven and capable of providing all of the thrust required for the set with one rotor inoperative, because this would increase the weight and complexity of the system. A coaxial rotor set with a particular diameter and solidity uses less total power to achieve a given thrust than a single rotor with the same diameter and solidity would use for the same thrust. Furthermore, the power and torque required by each rotor system of a coaxial set is approximately half the total power of the rotor system. This negates the advantage that coaxial rotor set designs, such as some of the designs implemented in some embodiments herein, have in the art felt of using coaxial rotors to reduce the power required. Oversizing a rotor system to enable VTOL flight with one rotor inoperable requires much more (if not less) installed power than a single rotor system.

[0018] What the prior art does not teach is a VTOL aircraft with fewer than five rotors, stacked or otherwise implemented in two main coaxial rotor stacks, that have low total rotating surface loads and are sized and dimensioned such that the aircraft can take off and land safely in a controlled manner even if the rotor system is completely inoperable. Such an approach contrasts with conventional wisdom which teaches that system redundancy requires a large number of rotors, typically five or more, to accommodate controlled VTOL flight after a complete failure of the rotor system, or that system redundancy would be weight prohibitive, and instead assume that the design of the vehicle assumes that at least one rotor or drive system is flight-critical, thereby increasing manufacturing and maintenance costs, even if feasible, by approximately 10 failures per flight hour. -9 This requires component-level reliability. [Means for solving the problem]

[0019] In some embodiments presented herein, the VTOL aircraft has four main rotors arranged in two main coaxial rotor stacks or otherwise implemented, designed and sized for redundancy of each rotor system. Such designs may replace an integral rotor system with a coaxial rotor set, each having two or more independently driven rotor systems. Other embodiments may utilize a non-coaxial rotor system, where the four main rotors are arranged in a quad configuration.

[0020] Unlike prior art coaxial rotors that are primarily designed for antitorque (no tail rotor), high edgewise advance ratios, or increased horsepower loading with respect to coaxial implementations, some embodiments described herein provide coaxial rotor sets designed for redundancy and safety. Specifically, each coaxial rotor set is designed such that each of the individual rotor systems in the set can be independently driven and provide all of the thrust required for the flight-critical coaxial rotor set for controlled VTOL flight, including takeoff.

[0021] In some coaxial embodiments, the rotor system is driven by a torque generating source (most likely a motor or engine), but the design may have a large excess installed power if each torque generating 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 motor or engine driving each rotor, are also sized so that a single rotor in the stack provides enough power to provide all of the thrust required for the entire stack. In the event that the aircraft is battery powered, the battery system architecture provides sufficient battery power for the motor or other torque generating device driving the operable rotor. Additionally, the aircraft has a method for balancing the yaw moment of the aircraft when one of the rotor systems in the coaxial set is inoperable. In some embodiments, the method may involve stopping one or more rotor systems of the other coaxial rotor sets, varying the speed and collective of the other coaxial rotor sets to vary their respective yaw moment contributions, applying differential rotor nacelle tilt angles to generate offsetting yaw moments, or other methods.

[0022] Some of the inventive concepts disclosed herein achieve the increased safety levels required for air transportation of people and cargo without the usual tradeoffs of high rotating surface loads or increased rotor count, and this approach allows for large diameter rotors without making those rotors and associated drivetrains safety critical.

[0023] Certain embodiments of the invention disclosed herein may provide apparatus, systems, and methods in which an electric vertical take-off and landing (eVTOL) aircraft is designed to carry at least 400 pounds using two coaxial rotor stacks, each stack being comprised of two variable pitch rotor systems axially aligned substantially about a common axis of rotation, with each variable pitch rotor being independently driven, preferably by one or more electric motors, but possibly 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 rigid (non-articulated) rotor. Moreover, the rotors are preferably tilt rotors configured to remain axially aligned relative to each other while tilting relative to the body. One or more of the 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 between the two variable pitch rotor systems in a coaxial stack is not fixed, and each variable pitch rotor system can rotate about a substantially common axis of rotation independently of the other variable pitch rotor system. To the extent the rotors tilt, both variable pitch rotor systems are configured to tilt together. The variable pitch rotor may include a first blade having a lbs / in at 30% of the rotor root radius. 2The flap stiffness of the first blade in a variable speed variable pitch rotor, such as the rotor taught in U.S. Pat. No. 6,641,365 (Karem), incorporated herein by reference, is at least 200 times the rotor diameter in feet to the fourth power. In some embodiments, the variable pitch rotor is configured to operate at an RPM level less than 80% or even less than 60% of the maximum rotor system RPM level.

[0024] Some embodiments of the invention disclosed herein are specifically contemplated for aircraft designed to carry at least 400 pounds, which is roughly the payload capacity required to carry two or more humans or an equivalent amount of other payload. VTOL aircraft are subject to a square-cube scaling law, where an aircraft scaled in size to a larger size by a factor increases weight by about the cube of the factor, while rotor area, wing area, structural spar depth (which provides strength and stiffness), and other similar parameters increase only about the square of the factor. The result is threefold. First, the required structural strength, stiffness, and aeroelastic margins become more difficult to maintain at larger scales. Second, the power required to hover, which already increases superlinearly with aircraft weight for a given rotor disk area, also increases superlinearly as the overall size of the vehicle increases. This means that the power margin required for propulsion redundancy increases undesirably with size. As a result, design techniques utilized in toys and smaller aircraft, especially those designed to carry the equivalent of less than two people, tend to fail in larger aircraft where the power requirements and design complexities are much higher. The inventive concepts contained herein may be particularly well suited to address the needs of aircraft capable of carrying more than 400 lbs. [Brief description of the drawings]

[0025] [Figure 1] FIG. 1 shows a prior art four-rotor embodiment of a vehicle with two to four flight-essential rotor systems. [Diagram 2] FIG. 1 illustrates a prior art aircraft. [Diagram 3] Diagram showing an aircraft with six stacks of two small co-rotating propellers. [Figure 4] A diagram showing Airspace Experience Technology's MOBI-ONE, a tilt-wing aircraft with eight small rotors arranged in four co-rotating coaxial stacks of two rotors each. [Figure 5A] A diagram showing Embraer's PULSE concept, a tilt-rotor aircraft with four rotors arranged in two tilting coaxial rotor stacks. [Figure 5B] An alternative view showing Embraer's PULSE concept, a tilt-rotor aircraft with four rotors arranged in two tilting coaxial rotor stacks. [Figure 6] Diagram showing a VTOL aircraft with four coaxial stacks of two rotors each for a total of eight rotors. [Figure 7] FIG. 1 illustrates a conventional aircraft. [Figure 8] A diagram showing Joby's six-rotor VTOL. [Figure 9] 1 is a perspective view of a preferred embodiment of a hovering VTOL aircraft in accordance with the inventive concepts herein; [Figure 10] FIG. 10 is a perspective view of the preferred embodiment of FIG. 9 shown in a transition state. [Figure 11] FIG. 10 is a perspective view of the preferred embodiment of FIG. 9 shown in a cruising position. [Figure 12] FIG. 9 illustrates a possible embodiment of a drive system 900 for a two rotor coaxial stack. [Figure 13] FIG. 9 illustrates a possible alternative embodiment of a drive system 900 for a two rotor coaxial stack. [Figure 14]FIG. 9 illustrates another possible alternative embodiment of a drive system 900 for a two rotor coaxial stack. [Figure 15] FIG. 9 illustrates a possible alternative embodiment of a drive system 900 for a two rotor coaxial stack. [Figure 16] FIG. 2 shows one possible embodiment of a reverser for the drive system. [Figure 17] FIG. 2 illustrates an embodiment in which the first wing is configured to pivot relative to the body of the aircraft. [Figure 18] FIG. 1 illustrates an embodiment of an aircraft including first and second coaxial rotor stacks as well as at least a first auxiliary rotor system. [Figure 19] FIG. 1 illustrates an embodiment of an aircraft configured with four controllable pitch rotors, where the aircraft is shown in a transition state. [Figure 20] FIG. 1 illustrates an embodiment of an aircraft configured with four controllable pitch rotors, the aircraft being shown in a hovering position. [Figure 21] FIG. 1 illustrates an embodiment of an aircraft configured with four controllable pitch rotors, with the aircraft shown in a cruising position. [Figure 22] FIG. 20 is a perspective view of the aircraft of FIG. [Figure 23] FIG. 1 illustrates a top view of an embodiment of an aircraft configured with four controllable pitch rotors. [Figure 24] A diagram showing possible reduction gear system, nacelle, 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). [Diagram 25] FIG. 1 illustrates a top view of an embodiment of an aircraft configured with four controllable pitch rotors. [Figure 26] FIG. 1 illustrates a possible implementation of an electronic flight control system and related systems. [Figure 27] FIG. 2 illustrates another possible implementation of the electronic flight control system and related systems. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0026] FIG. 9 is a perspective view of a preferred embodiment of a hovering VTOL aircraft. The aircraft has a body 1101, an inboard wing 1102, a tilting nacelle 1106, a tilting outboard wing 1107, a first coaxial rotor stack 1103, and a first empennage surface 1108. The outboard wing 1107 tilts with the tilting nacelle 1106 to reduce download forces from the rotor in hover mode. The first coaxial rotor stack 1103 includes a first variable pitch rotor system 1104 and a second variable pitch rotor system 1105 that are rotatable about a substantially common axis of rotation, i.e., the rotor axis of rotation, with each of the rotor systems independently driven by at least one torque source. The first and second variable pitch rotors may be configured to co-rotate or counter-rotate.

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

[0028] The first variable pitch rotor system 1104 includes rotor blades 1109. The rotor blades are of the rigid hingeless variety, such as those described in U.S. Pat. No. 6,641,365 (Karem). Each variable pitch rotor system independently provides thrust and force moment controlled by rotating the blades about a feather axis, such as that of the rotor system described in U.S. Pat. No. 10,351,235 (Karem), which is incorporated herein by reference, and 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. Pat. No. 6,641,365 to trim the rotor forces and moments in accordance with the electronic flight control system and ensure that the rotors operate at optimal speed for optimal blade lift distribution, and therefore optimal efficiency and minimum required rotor torque applied by the torque generating device. The inner blade 1102 transfers loads from the coaxial rotor stack 1103 and the outer blade 1107 to the body 1101 .

[0029] Each of the first and second controllable pitch rotor systems of each of the coaxial rotor stacks can provide sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if at most one of the controllable pitch rotor systems of each of the coaxial rotor stacks is inoperative. This requires that the total thrust required for each coaxial stack, including the inefficiency of the additional rotor due to the inoperative rotor in the wake, can be provided by a single rotor. To calculate the total required thrust, the maximum total weight of the aircraft is considered, and then computational fluid dynamics (CFD) methods are used to calculate the additional thrust required to overcome the wing, body, nacelle, tail, and downwind loads due to the inoperative rotor that may affect the rotor wake. In addition, additional thrust margin is added for controllability and maneuverability with one rotor inoperative, which in the preferred embodiment may be greater than 15%. CFD methods are then used to calculate the rotor input torque and power required to achieve the required thrust, including margins, at the desired operating conditions, including air density and rate of climb, if one rotor is inoperable. This required rotor torque is then used to size the torque and power ratings for each element of the powertrain, including the reduction system, the torque generating device, and the power delivery system (including the battery in an embodiment where the torque generating device is an electric motor), with the power and torque ratings of each powertrain element further accounting for efficiency losses in each element of the powertrain system.

[0030] Each of the first and second variable pitch rotor systems 1104 and 1105 in the coaxial rotor stack 1103 may be sized such that each of the first and second variable pitch rotor systems can provide all of the thrust required by the electronic flight control system for the entire coaxial rotor stack even if one rotor is inoperative. To achieve this, the electronic control system may request increased RPM and power for the operable variable pitch rotor. The increased RPM and power for the required thrust margin is calculated by computational fluid dynamics (CFD) methods for the specific rotor configuration, including the presence of the inoperative rotor, when sizing the rotor with results validated by physical rig testing of the rotor, nacelle, and adjacent blade surface combination. In flight, the required RPM, blade pitch, and power are requested by the electronic flight control system based on measured feedback of the aircraft's linear and angular velocity and acceleration about all three axes. The torque generating devices and reduction systems (shown in Figures 11, 12, and 13) are similarly sized with sufficient margin in the event one rotor is inoperable. If the aircraft is battery powered, the batteries and distribution system are sized to provide sufficient power for the operable rotors in each scenario in which a rotor becomes inoperable. Such sizing is performed at maximum total takeoff weight at the highest pressure altitude desired for operation with a sufficient vertical climb rate, typically greater than 100 feet per minute, and additional margins for control power, crosswinds, and trim penalties as dictated by the desired flight envelope.

[0031] The body 1101 is designed to carry a payload such as passengers, baggage, or cargo, and houses various systems including landing gear in a configuration such as that described in detail in U.S. Patent No. 10,351,235 (Karem), incorporated herein by reference. In a preferred embodiment, deviation of the vehicle's center of gravity from the center of the coaxial rotor stack is controlled by mast moments from a rigid variable pitch rotor system, which necessarily requires rotational control of the rotors.

[0032] In one exemplary embodiment, the aircraft of FIG. 9 has a maximum gross takeoff weight of 6000 lb. The aircraft has one cockpit and four passenger seats, and a baggage compartment. The aircraft body is 33 ft long, 5.0 ft wide at its widest point, and 5.5 ft high at its widest point. The aircraft wingspan is 45 ft, and the spanwise distance between the centers of the two coaxial rotor stacks is 30 ft. The port and port coaxial rotor stack controllable pitch rotor system is 64.3818 m. 2 (693ft 2 ) non-overlapping rotor disk area with a diameter of 6.4008 m (21 ft), based on a completely non-overlapping area the maximum torque is 42.4771 kg / m 2 (8.7lb / ft 2 ), based on a completely overlapping disk area it is 20.9944 kg / m 2 (4.3lb / ft 2) rotating surface load. Each variable pitch rotor system is driven by four motors through a single speed reduction system, and in hovering during nominal operating conditions, the motor input speed is 10,000 RPM and the variable pitch rotor system output speed is 400 RPM. However, any number of electric motors, e.g., one, two, or three, may be configured to drive each variable pitch rotor system through the reduction system. Each motor preferably has a maximum continuous power rating of 90 kW and a maximum emergency power rating of 110 kW, although the vehicle will typically use less than half of its maximum continuous power 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] In the event that one variable pitch rotor system becomes inoperable, the electronic control system increases the speed of the remaining operable rotor to 533 RPM and increases the blade pitch to provide the same thrust as in nominal operating conditions while maintaining rotor operation near the rotor's peak efficient thrust coefficient. Importantly, oversizing the motors in some embodiments described herein allows the vehicle to maintain positive performance margins in both vertical climb in ground effect (at least 100 ft / min climb) and control power (e.g., at least 25% control power margin for a vehicle operating in an urban environment) at a pressure altitude of at least 5000 ft. In a rotor system failure condition, the vehicles of some embodiments described herein allow the pilot to maintain controlled flight and transition to wing flight where much less power is required, and also allow the pilot to take off vertically and climb in ground effect if required, such as in an aborted landing scenario. The same applies if the battery pack fails while all rotors are operating at nominal speed, and you can also oversize the batteries to allow for a combined battery and rotor drive failure.

[0034] 10 and 11 are perspective views of the preferred embodiment in transition and cruise, respectively, consistent with the aircraft shown in FIG. 9. As the tilt nacelle 1106 tilts the coaxial rotor stack 1103 about tilt axis 1113 from a hover position with a vertical thrust vector 1111 (shown in FIG. 9) to a cruise position with a horizontal thrust vector 1112 (shown in FIG. 11), the outer wing 1107 tilts from a vertical position to a horizontal position at a similar angle of attack as the inner wing 1102. The outer wing increases the wing aspect ratio of the overall vehicle with reduced cruise 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 preferred level of safe and efficient flight during hover, transition, and cruise, the aircraft contemplated herein may be designed in accordance with the aerodynamic design teachings of U.S. Pat. No. 10,351,235 (Karem), which is incorporated herein in its entirety. In such an embodiment, 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 hover mode to cruise mode at low airspeeds. Such aircraft may advantageously use slotted wing flaps as taught in U.S. Provisional Patent Application No. 62 / 757,680 (Karem), which is incorporated herein in its entirety, to reduce the stall speed of the vehicle while providing an efficient wing area for higher speed cruise. In some embodiments, the vehicle is designed to be efficient at cruising with a lift to drag ratio of at least 9, and preferably at least 13, at a cruise speed of 130 KTAS.

[0036] 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 an aft rotor system 902 rotating about a common axis 903. FIG. 12 shows a dual common assembly and FIG. 13 shows a back-to-back motor-reduction system arrangement. The rotors can co-rotate by the ability to stationary either the forward or aft rotor when it becomes unavailable while the available rotor continues to rotate. The forward rotor hub 904 and the aft rotor hub 905 are driven by the forward motor 906 or the aft motor 907 through a reduction gearing 908. The hubs 904 and 905 rotate on bearings 909 that are coaxially mounted on a structural nacelle extension 910. The collective and cyclic pitch of the vanes 911, with vane shanks 912 located in feather bearings 913, may be controlled by rotary actuators 914 located in hubs 904 and 905, and electronic control of both motor speed and vane pitch is performed by a control and driver module 915. The drive system is lubricated and cooled by oil contained in a tank 916, and independent spinner assemblies 917 and 918 surround these hub systems.

[0037] In one embodiment, such as the embodiment shown in FIG. 14, the drive system 900 can be configured such that the forward motor 906 and the aft motor 907 are torque-connected by a connecting shaft 919 that allows for power sharing between the forward variable pitch rotor system and the aft variable pitch rotor system. In such an embodiment, the drive system 900 is configured such that the aft motor 907 and the forward motor 906 can drive either the forward rotor hub 904 or the aft rotor hub 905, or both the rotor hub 904 and the rotor hub 905. In such a configuration, for example, if the aft reduction system 908 becomes jammed, the forward hub can be driven by the forward motor and the aft motor without the aft hub being driven. Similarly, the aft hub can be driven by the forward motor and the aft motor without the forward hub being driven. Although not visible in Figure 24, a clutch such as directional clutch 920a or 920b can be configured such that driving the forward and reverse motors in a first direction drives a first variable pitch rotor system and driving the forward and reverse motors in a second direction drives a second variable pitch rotor system. Such an embodiment is preferred because the motors can be sized such that four motors are used to drive a single operable rotor system, rather than two sets of motors each driving their own motor, thus reducing the installed power requirements to drive only one of the variable pitch rotor systems.

[0038] As shown in FIG. 15, some embodiments may implement motors torque-connected to both rotors in a coaxial stack. Motors 906 and 907 may be interposed between sections of a connecting shaft 919. Directional clutches 920a and 920b may be interposed between each end of the connecting shaft 919 and the respective reduction systems. For example, in a counter-rotating embodiment, motor 906 may be configured to drive the connecting shaft 919 in a first direction. The forward directional clutch 920a may engage and drive the forward reduction system and thus the forward rotor. During normal operating conditions, the rearward directional clutch 920b connected to the same forward motor 906 is configured to be disengaged. In the event of a failure of the forward rotor, the forward motor 906 may reverse its direction of rotation, thereby disengaging the forward directional clutch 920a and engaging the rearward directional clutch 920b. Thus, all motors may be configured to generate thrust even if one rotor becomes inoperable. Similarly, the rear motor 907 may be configured to drive the rear rotor during nominal operating conditions and drive the forward rotor if the rear rotor becomes inoperable. The directional clutches 920a and 920b may be configured to provide desired nominal and failure mode operating characteristics. Although only three motors are visible in FIG. 15, it should be understood that there could be six motors (two additional rear motors 907 and one additional forward motor 906). Other embodiments could include any number of forward and rear motors, e.g., two or four. The motors are shown intermediate the forward and aft rotor hubs, but could be located anywhere. Other implementations could include alternative methods of decoupling the connecting shaft from the inoperable rotor and allowing it to drive the still-operable rotor, including the shear shaft.

[0039] Co-rotating embodiments may also include a reverser, such as the reverser 1501 shown in FIG. 16, configured to allow the rear motor 907 and the forward motor 906 to drive either the forward hub 904 or the rear rotor hub 905. The reverser 1501 may have a pinion gear 1502, a reverser gear 1503, and a ring gear 1504 driven by an associated motor. The electric motor may be configured to drive the pinion gear 1502 such that the ring gear 1504 drives the respective hub in the opposite direction than if the reverser were not interposed.

[0040] 17, the first wing 2202 is configured to pivot relative to the body 2201. In some embodiments, the first wing 2202 may comprise a left wing 2203 and a right wing 2204, each configured to pivot independently about a substantially spanwise axis.

[0041] In another embodiment shown in FIG. 18, the aircraft includes at least a first auxiliary rotor system 1115, such as an auxiliary rotor disclosed in U.S. Patent Application Publication No. 2018 / 334251 (Karem), which is incorporated herein by reference, in addition to the first and second coaxial rotor stacks. The aircraft may also include a second auxiliary rotor system 1115. In such an embodiment, the first and second variable pitch rotor systems of each of the coaxial rotor sets and each of the first and second auxiliary rotor systems may cooperate to provide sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if either at most one of the variable pitch rotor systems of each of the coaxial rotor stacks or at most one of the auxiliary rotor systems is inoperative. In one embodiment, the first auxiliary may have a rotor having a diameter of 50% or less of the diameter of the first variable pitch rotor system. However, in some embodiments, the auxiliary rotor may be more than 50%. Additionally, each auxiliary rotor system may include a controllable pitch rotor.

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

[0043] Four rotors may be the preferred rotor count because it may allow for VTOL operation without all four rotors being operational, but also allows for larger rotors. A fail-operational four-rotor tilt rotor that can perform VTOL flight without all rotors in use may have unexpected advantages.

[0044] 20, an aircraft is configured with four variable pitch rotors: 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 inboard wing 1402, a tilting nacelle 1406, a tilting outboard 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 FIG. 23, the aircraft is configured such that the center of gravity 1409 is located approximately at the intersection of a first diagonal 1412 and a second diagonal 1413, where the first and second diagonals connect the centers of thrust 1403 of the controllable pitch rotors of the first pair of controllable pitch rotors 1404 and the second pair of controllable 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. One skilled in the art would understand that the center of gravity 1409 moves within a small envelope based on passenger and payload loading, etc. Thus, the location of the rotors relative to the aircraft's center of gravity is configured such that the aircraft can perform controlled vertical takeoff and landing even if at most one of the controllable 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 rotors. The left variable pitch rotors of the body may co-rotate. Additionally, the right variable pitch rotors may also co-rotate. The right variable pitch rotor may counter-rotate with the left variable pitch rotor. The aircraft may have a second wing 1402. Preferably, the at least one torque generating source comprises an electric motor, but may alternatively comprise a fuel consuming engine.

[0045] In one particularly preferred embodiment, the aircraft of FIG. 20 may have a maximum gross takeoff weight of 6,000 pounds and may be configured to carry at least 400 pounds. The aircraft has a cockpit and four passenger seats and a baggage compartment. The aircraft body is 35 feet long, 5.0 feet wide at its widest point and 5.5 feet high at its widest point. The aircraft wingspan is 49 feet and the spanwise distance between the centers of the two forward controllable pitch rotors is 29 feet. The controllable pitch rotors are 49 feet wide and 116 feet high. 2 (1,257ft 2 ) rotor disk area, diameter 6.096 m (20 ft), 23.2892 kg / m 2 (4.77 lb / ft 2 ) rolling surface load. Each variable pitch rotor is driven by four motors through 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, for example one, two, or three, may be configured to drive each variable pitch rotor through the reduction system. Each motor has a maximum continuous power rating of 70 kW and a maximum emergency power rating of 90 kW, although the vehicle will typically use less than 40% of its maximum continuous power rating in hovering during 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 an embodiment, the controllable pitch rotors may be sized so that no single rotor is flight-critical. -9 For flight in urban environments where a degree of system reliability is desired, the embodiment shown in FIG. 20 may allow that level of safety to be achieved at the aircraft level without requiring the same subsystem or component reliability.

[0047] Figure 21 shows an alternative view of the same embodiment shown in Figure 20. The center of gravity 1409 is shown. The wing section 1402 is also shown. Figure 22 shows an alternative view of the same embodiment shown in Figure 20. This embodiment is shown in wing flight mode. Thrust vectors 1405 are shown.

[0048] In one embodiment shown in FIG. 23, controlled vertical takeoff and landing is achieved by reducing power to the controllable pitch rotor that is approximately across an envelope that includes the center of gravity 1409 from the inoperable controllable pitch rotor. For example, if the forward left controllable pitch rotor 1404 becomes inoperable due to a failure, the electronic flight control system 1411 may reduce or cut power to the aft right controllable pitch rotor 1404. Alternatively, the mast moment of the remaining controllable pitch rotor may be used to correct the imbalance created by the inoperable rotor. A combination of mast moment control and thrust control may also be used to achieve controlled vertical takeoff and landing with less than all four controllable pitch rotors. The controllable pitch rotors 1404 and the associated torque providing sources are configured to provide sufficient thrust to achieve vertical takeoff and landing with one or more controllable pitch rotors inoperable.

[0049] A configuration of the reduction system, nacelle, and motors such as that shown in FIG. 24 may be used. A drive train and nacelle configuration taught in U.S. Pat. No. 10,351,235 (Karem), incorporated herein by reference, may also be used. In FIG. 24, three motors 1905 drive the ring gear 1912. Such a system provides good additional redundancy at the motor level. In the configuration shown in FIG. 24, each variable pitch rotor is driven by three motors 1905, although other numbers of motors are also contemplated. For example, two motors or four motors per rotor are also contemplated. However, other drive train configurations may also be used with this embodiment.

[0050] A particularly preferred embodiment comprises an electronic flight control system 1411 that is configured to respond to faults without requiring discontinuous mode switching based on a diagnosis of the fault. Such an embodiment eliminates the risk of misdiagnosing the condition of the aircraft. One embodiment achieves this objective by partitioning the functions of the electronic flight control system 1411 such that subsystems in the electronic flight control system handle fault conditions by architectural design rather than by real-time decision making.

[0051] In one embodiment shown in FIG. 25 and FIG. 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 with the vehicle 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 that make up each variable pitch rotor pair are on opposite sides of approximately the center of gravity 1409. It should be understood that the two variable pitch rotors that comprise a variable pitch rotor pair may be on opposite sides of an envelope that encompasses the center of gravity 1409. Similarly, the third variable pitch rotor and the fourth variable pitch rotor 1404 are on opposite sides of an envelope that includes the center of gravity.

[0052] As shown in FIG. 25, a horizontal plane is defined by the roll axis 1422 and pitch axis 1416 of the aircraft. The x-axis 1420 of the variable pitch rotor pair runs between the centers of the variable pitch rotors of each variable pitch rotor pair 1414 through an envelope encompassing the center of gravity 1409. The coordinate frame of the first variable pitch rotor pair has an x-axis 1420 that intersects with the centers of the variable pitch rotors of the first variable pitch rotor pair. The coordinate system of the first variable pitch rotor pair further includes a y-axis 1415 of the corresponding variable pitch rotor pair that is orthogonal to the x-axis and runs parallel to the horizontal plane. The coordinate system of the first variable pitch rotor further includes a z-axis of the first variable pitch rotor pair that is orthogonal to both the x-axis and y-axis of the variable pitch rotors and may correspond to altitude. The coordinate system of the first variable pitch rotor further comprises a first moment axis 1419 about the x-axis of the variable pitch rotor pair, and a second moment axis 1423 about the y-axis 1415 of the variable pitch rotor pair. The coordinate frame of the second variable pitch rotor pair has an x-axis 1420 that intersects the center of the variable pitch rotors comprising the second variable pitch rotor pair 1414. The coordinate system of the second variable pitch rotor pair further comprises a y-axis 1415 of the corresponding variable pitch rotor pair that is orthogonal to the x-axis 1420 and runs parallel to the horizontal plane. The coordinate system of the second variable pitch rotor further comprises a z-axis of the first variable pitch rotor pair that is orthogonal to both the x-axis and y-axis of the variable pitch rotors, and may correspond to altitude. The coordinate system of the second variable pitch rotor further comprises a first moment axis 1419 about the x-axis of the variable pitch rotor pair and a second moment axis 1423 about 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 about the y-axis 1415 of each variable pitch rotor pair. The vertical thrust due to each variable pitch rotor pair 1414 is the sum of the thrust due to each variable pitch rotor 1404 in each variable pitch rotor pair 1414. The moment about the y-axis of each variable pitch rotor pair is the thrust difference between the two rotors that comprise each variable pitch rotor pair 1414 multiplied by the distance between the centers of the respective variable pitch rotors.

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

[0055] The electronic flight control system 1411 may receive vehicle dynamics values ​​from the vehicle dynamics sensors 1421. The vehicle dynamics sensors 1421 may comprise one or more of a GPS, magnetometer, IMU, or other known sensors. The electronic flight control system may use values ​​from the vehicle dynamics sensors 1421 to calculate vehicle dynamics error values. The vehicle dynamics error values ​​are transformed 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. The system attitude input is added to or subtracted from a corresponding system altitude input. The combined input is then sent to a vehicle dynamics control device. 26, these signals are shown as being sent to a vehicle dynamics control device associated with each rotor, such as a rotor blade pitch actuator and motor, but it should be understood that these signals may be sent to other vehicle dynamics control devices, which may include an electric motor coupled to one of the variable pitch rotors, an actuator configured to control the blade pitch of the 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 FIG. 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, but as long as the other variable pitch rotor in the variable pitch rotor pair 1414 is functional, the variable pitch rotor pair retains its ability to generate a combination of thrust and moment. In the event of a failure that renders one rotor inoperable, the variable pitch rotor pair with the failure continues to adjust the vehicle attitude but is unable to contribute to adjusting the vehicle altitude. The other variable pitch rotor pair 1414, consisting of two normally functioning variable pitch rotors, assumes responsibility for adjusting altitude.

[0057] The electronic flight control system 1411 may include a look-up table of thrust coefficients for collective that the electronic flight control system may use to determine preferred RPM and collective pitch settings for a sensed flight condition. Configuring the electronic flight control system 1411 with the look-up table allows an ideal RPM and collective pitch to be used 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 altitude and attitude of the vehicle. One advantage of such an embodiment is that the electronic flight control system can be configured to provide continued smooth control when the aircraft experiences component failure.

[0059] In the event that one rotor becomes inoperable, the electronic flight control system 1411 may reduce or cut power to one of the operable control pitch rotors. The electronic flight control system 1411 may also vary the mast moment of the operable control pitch rotor to provide a combination of forces such that the aircraft may perform vertical takeoff and landing with not all of the control pitch rotors 1404 in operation. The electronic flight control system 1411 may alternatively reduce power to at least one control pitch rotor and vary the pitch moment of the operable control pitch rotor 1404 to provide a desired combination of forces.

[0060] Similarly, an embodiment of the aircraft contemplated herein may be configured to achieve VTOL flight with two of the controllable pitch rotors inoperable if the inoperable controllable pitch rotors are located at diagonal corners. For example, in FIG. 23, if the forward left controllable pitch rotor 1404 and the aft right controllable pitch rotor 1404 are both inoperable, VTOL flight may be achieved using only the forward right controllable pitch rotor 1404 and the aft left controllable pitch rotor 1404.

[0061] It should be noted that any reference to an electronic flight control system, or control and driver module, should be read as including any suitable combination of computing devices, such as servers, interfaces, systems, databases, agents, peers, engines, controllers, or other types of computing devices acting individually or collectively. These computing devices may also comprise processors configured to execute software instructions stored in a tangible, non-transitory computer-readable storage medium (e.g., hard drive, solid state drive, RAM, flash, ROM, etc.). The software instructions preferably configure the computing device to provide the roles, responsibilities, or other functions described above with respect to the disclosed apparatus. In some embodiments, the 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 services APIs, known financial transaction protocols, or other methods of electronic information exchange. The data exchange is preferably performed over a packet-switched network, the Internet, a LAN, a WAN, a VPN, or other type of packet-switched network. Aspects of the electronic flight control system may be located anywhere on the aircraft where the actuators are located, or may be located anywhere else, such as in a ground control center, on another aircraft, or even as part of the actuators themselves.

[0062] Aspects of the electronic flight control system may be located anywhere on the aircraft, or may be located anywhere else, such as in a ground control center, on another aircraft, etc. Additionally, in some embodiments, the electronic flight control system and the control and driver modules may be implemented in distinct units or may be combined into one unit.

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

[0064] Although several embodiments have been described herein, they are to be understood in all respects as illustrative and not limiting, and it is to be understood that it is contemplated that aspects of various embodiments may be combined with aspects of other embodiments.

Claims

1. 1. An aircraft capable of carrying a payload of at least 400 pounds, The main body, a first variable pitch rotor, a second variable pitch rotor, a third variable pitch rotor, and a fourth variable pitch rotor, each of said rotors independently driven by at least one torque source; each of the first, second, third, and fourth variable pitch rotors is a variable speed rotor; 1. An aircraft, wherein 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 such that the aircraft can perform controlled vertical takeoff and landing even when only three of the variable pitch rotors are operational.

2. The aircraft of claim 1 , further comprising a first wing mechanically coupled to the body.

3. The aircraft of claim 2 , further comprising a second wing mechanically coupled to the body.

4. The aircraft of claim 2 , wherein the first wing is configured to pivot relative to the body.

5. 2. The aircraft of claim 1, wherein the first and second controllable pitch rotors are on opposite sides of an envelope that includes a center of gravity, and the first and second controllable pitch rotors are cooperatively capable of providing all of the thrust required to enable the aircraft to perform controlled vertical takeoff and landing.

6. 10. The aircraft of claim 1, further comprising a tilt mechanism configured to tilt the first variable pitch rotor at least 80 degrees.

7. 2. The aircraft of claim 1, wherein the first and second variable pitch rotors are on a port side of the body and the third and fourth variable pitch rotors are on a starboard side of the body, and the first and second variable pitch rotors are configured to rotate in a first direction.

8. 8. The aircraft of claim 7, wherein the third and fourth variable pitch rotors are configured to rotate in a second direction, the second direction being opposite the first direction.

9. The aircraft of claim 1 , wherein the first variable pitch rotor is configured to provide rotor rotation control.

10. 10. The aircraft of claim 9, wherein the first variable pitch rotor comprises rigid blades and a hub configured to apply a mast moment to the aircraft.

11. 11. The aircraft of claim 10, wherein the first variable pitch rotor comprises rotor blades with individual blade control.

12. 10. The aircraft of claim 1, wherein the aircraft comprises four or fewer controllable pitch rotors.

13. 2. The aircraft of claim 1, wherein the first torque source is a first electric motor.

14. 14. The aircraft of claim 13, further comprising a reduction gear system operatively coupling the first variable pitch rotor and the first electric motor.

15. 15. The aircraft of claim 14, wherein the first variable pitch rotor is driven by each of the first electric motor and at least a second electric motor through the reduction system.

16. 2. The aircraft of claim 1, wherein the first variable pitch rotor is configured to operate at an RPM level that is less than 60% of a maximum rotor RPM level.

17. 10. The aircraft of claim 1, wherein an 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.

18. 2. The aircraft of claim 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 discontinuous switching of states.

19. 1. A method of operating a VTOL aircraft, comprising: receiving a command for a vehicle attitude; requesting thrust from one of four variable pitch rotors; receiving a vehicle dynamics signal from a vehicle dynamics sensor; determining, using the received vehicle dynamics signals, that the thrust required for the variable pitch rotor is insufficient to achieve a required vehicle attitude; increasing the thrust required from each of the controllable pitch rotors until sufficient thrust is generated to perform a controlled vertical takeoff and landing; A method comprising:

20. 20. The method of claim 19, comprising the additional step of generating sufficient thrust so that the aircraft performs controlled vertical takeoff and landing even when only three of the controllable pitch rotors are operational.

21. 1. An aircraft capable of carrying a payload of at least 400 pounds, The main body, first and second coaxial rotor stacks, each having a first variable pitch rotor system and a second variable pitch rotor system rotatable substantially about a common axis of rotation, each of said rotor systems being independently driven by at least a first torque generating source; 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; wherein each of the first and second variable pitch rotor systems of each of the coaxial rotor stacks provides sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if at most one of the variable pitch rotor systems of each of the coaxial rotor stacks is inoperative.

22. 22. The aircraft of claim 21, further comprising a wing mechanically coupled to the body.

23. 23. The aircraft of claim 22, wherein the wing is configured to pivot relative to the body.

24. 24. The aircraft of claim 23, wherein the wing is divided into a left wing and a starboard wing, each of the left and starboard wings configured to pivot independently about a substantially spanwise axis.

25. 22. The aircraft of claim 21, further comprising a tilt mechanism configured to tilt the first coaxial rotor stack at least 80 degrees.

26. 22. The aircraft of claim 21, wherein the first and second variable pitch rotor systems of the first coaxial rotor stack are configured to rotate in the same direction.

27. 22. The aircraft of claim 21, wherein the first and second variable pitch rotor systems of the first coaxial rotor stack are configured to rotate in opposite directions.

28. 22. The aircraft of claim 21, wherein the first variable pitch rotor system of the first coaxial rotor stack is configured to provide rotor rotation control.

29. 30. The aircraft of claim 28, wherein the first variable pitch rotor system of the first coaxial rotor stack comprises rigid blades and a hub configured to apply a mast moment to the aircraft.

30. 30. The aircraft of claim 28, wherein the first variable pitch rotor system of the first coaxial rotor stack comprises rotor blades with individual blade control.

31. 22. The aircraft of claim 21, wherein the first torque source is a fuel consuming engine.

32. 32. The aircraft of claim 31, wherein the first torque source is a first electric motor.

33. 33. The aircraft of claim 32, further comprising a reduction gear system operatively coupling the first variable pitch rotor system and the first electric motor.

34. 34. The aircraft of claim 33, wherein the first variable pitch rotor system of the first coaxial rotor stack is driven by each of the first electric motor and at least a second electric motor through the reduction system.

35. 32. The aircraft of claim 31, further comprising a first mechanical clutch system, wherein at least the first torque generating 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.

36. 32. The aircraft of claim 31, wherein the first and second variable pitch rotor systems of each of the first and second coaxial rotor stacks, together with each of the torque generating sources and together with an electronic control system, are sufficiently sized and configured to enable controlled vertical takeoff and landing without the aid of an additional lift generating source even if at most one of the variable pitch rotor systems of each of the coaxial rotor stacks is inoperative.

37. 32. The aircraft of claim 31, wherein the first variable pitch rotor system is configured to operate at an RPM level that is less than 80% of a maximum rotor system RPM level.

38. 40. The aircraft of claim 37, wherein either the first or second variable pitch rotors are configured to operate at an RPM level that is less than 60% of a maximum rotor system RPM level.

39. 1. An aircraft capable of carrying a payload of at least 400 pounds, The main body, first and second coaxial rotor stacks, each having a first variable pitch rotor system and a second variable pitch rotor system rotatable substantially about a common axis of rotation, each of said rotor systems being independently driven by at least one torque generating 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; at least a first auxiliary rotor system not contained in the first or second coaxial rotor stacks; wherein each of the first and second variable pitch rotor systems of each of the coaxial rotor sets and the first auxiliary rotor system cooperate to provide sufficient thrust to enable the aircraft to perform controlled vertical takeoff and landing even if either at most one of the variable pitch rotor systems of each of the coaxial rotor stacks or at most one of the auxiliary rotor systems is inoperative.

40. 40. The aircraft of claim 39, further comprising a second auxiliary rotor.

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