阅读说明：本技术 使用大型变速倾转旋翼的evtol飞行器 (EVTOL aircraft using large variable-speed tiltrotors ) 是由 阿贝·凯瑞姆 威廉姆·马丁·韦德 于 2018-05-21 设计创作，主要内容包括：设想了用于电动垂直起降(eVTOL)飞行器的装置,系统和方法。这样的飞行器经过工程设计,可以使用几个(例如2-4个)旋翼,通常是变速刚性(非铰接式)旋翼,安全地运载至少500磅(约227公斤)物体。可以设想,一个或多个旋翼在旋翼飞行期间(例如,垂直起飞,悬停等)产生大量的升力(例如70％),并且在机翼飞行中倾斜以提供向前推力。旋翼优选采用独立桨叶控制,并由电池供电。该交通工具优选以自动驾驶或无人驾驶模式飞行,并且具有较小的足迹(例如,直径小于45英尺)。(Apparatuses, systems, and methods for an electric vertical take-off and landing (eVTOL) aircraft are contemplated. Such aircraft are engineered to use several (e.g., 2-4) rotors, typically variable speed rigid (non-articulated) rotors, to safely carry at least 500 pounds (about 227 kilograms) of objects. It is contemplated that one or more rotors generate a substantial amount of lift (e.g., 70%) during rotor flight (e.g., vertical takeoff, hover, etc.) and tilt in wing flight to provide forward thrust. The rotor is preferably controlled by independent blades and is battery powered. The vehicle preferably flies in an autonomous or unmanned mode and has a small footprint (e.g., less than 45 feet in diameter).)

1. An electric VTOL aerial vehicle capable of carrying at least 500 pounds of payload, comprising:

a machine body

A wing mechanically coupled to the fuselage;

at least first and second main-shift rigid rotors configured to provide lift to the fuselage, wherein each of the first and second rotors is open and not embedded in the wing;

each of the first and second main rotors being driven by at least one electric motor, and the first and second main rotors being sized and dimensioned to collectively provide at least 70% lift at vertical takeoff; and

the at least one motor is powered by at least a first power source.

2. The aircraft of claim 1 further comprising a power plant/rotor assembly configured to tilt the first rotor by at least 80 °.

3. The aircraft of claim 1 wherein each of the first and second main rotors is configured to apply a moment to control the pitch of the aircraft in VTOL and wing cruise flight.

4. The aircraft of claim 3 wherein the first and second main rotors comprise a plurality of blades and hubs configured to provide a moment at least equal to 6% of the rotor maximum lift multiplied by the rotor radius.

5. The aircraft of claim 1 wherein each of the first and second main rotors is configured to provide a paddle load below 10psf and to provide a hover power load above 8 lb/HP.

6. The aircraft of claim 1 wherein the wing is configured to provide a wing load of no greater than 40psf, and wing-on-board stall of no greater than 90 KIAS.

7. The aircraft of claim 6 wherein the wing is further configured to provide a flight speed margin of no less than 20KIAS in a transition from rotor flight to wing flight.

8. The aircraft of claim 6 wherein the wing is further configured to provide a wing cruise lift/drag ratio to the aircraft of no less than 10.

9. The aircraft of claim 6 wherein the wing has a wing tip portion with a dihedral that actuates to move the wing tip to between 20-90 degrees.

10. The aircraft of claim 6 wherein the wing has an actuated slotted flap that provides a cross-sectional stall lift coefficient of at least 2.0.

11. The aircraft of claim 10 wherein the slotted flap has a deflection of at least 5 degrees upward and downward to act as an aileron for aircraft roll control.

12. The aircraft of claim 1 further comprising a first auxiliary rotor that is no more than 50% of the disc area of the first main rotor.

13. The aircraft of claim 12 further comprising a second auxiliary rotor that is no more than 50% of the disc area of the first main rotor.

14. The aircraft of claim 12 wherein the first auxiliary rotor provides a maximum aircraft pitching moment during rotor operation that is no greater than the collective total aircraft pitching moment capability of the first and second main rotors.

15. The aircraft of claim 1, further comprising a tail raised surface having an area between 10% and 100%, inclusive, of the wing area.

16. The aircraft of claim 1 further comprising a duckbill lift surface having an area between 10% and 100%, inclusive, of the wing area.

17. The aircraft of claim 1 wherein the fuselage has a passenger compartment with at least one seat configured for a person to sit on.

18. The aircraft of claim 1, wherein the at least first power source is at least partially disposed in the wing.

19. The aircraft of claim 1, wherein the at least first power source is at least partially disposed in the nacelle.

20. The aircraft of claim 1, further comprising a landing gear extending from at least one of the fuselage and the wing.

21. The aircraft of claim 1, further comprising electronics configured to fly the aircraft without an onboard pilot.

22. The aircraft of claim 1 further comprising an independent blade control system for each of the first and second main rotors, the independent blade control systems having a differential collective pitch between blades on one rotor such that rotor thrust remains substantially constant while increasing shaft torque above a desired torque without differential collective.

23. The aircraft of claim 1 wherein each blade on each of the first and second main rotors has an actuator internal to the blade, the actuator being axially mounted to the pitch axis.

24. The aircraft of claim 1 further comprising a rotating hub, a hub bearing, a gearbox, and a motor mounting fixture, which together are configured as an integrated rotor drive system.

25. The aircraft of claim 1 wherein said first main rotor has a rotor hub and at least two blades radially connected to said hub, each blade having a root near said hub and a tip distal to said hub, wherein the weight of each blade in pounds does not exceed the product of rotor diameter in cubic feet multiplied by 0.004.

26. The aircraft of claim 25 wherein the weight per blade in pounds does not exceed the product of rotor diameter in cubic feet multiplied by 0.004.

27. The aircraft of claim 25 wherein the flap stiffness of each blade at 30% of the rotor radius is in lbs-in measured from the rotor center of rotation²Not less than the fourth power of the product of rotor diameter in feet times 200.

28. The aircraft of claim 25 wherein the flap stiffness of each blade at 30% of the rotor radius is in lbs-in measured from the rotor center of rotation²Not less than the fourth power of the product of rotor diameter in feet times 200.

29. A method of engineering an aircraft, comprising:

engineering-wise, the design has 2, 3 or 4 tilt rotors, wherein at least one tilt rotor is powered by a first motor and the rotor disc load of each rotor is below 7lb/Ft 2); and

including in the design a wing with a wing load below 24lb/Ft ^ 2; and

configuring a flight weight of the aircraft to be between 500 pounds and 10000 pounds.

30. The method of claim 29, further comprising including a fuel-powered engine in the design that indirectly powers the first electric machine.

31. The method of claim 29 further including a battery in the design, the battery supplying at least some power to the first motor.

32. The method of claim 29 further including a battery in said design, said battery providing all of the power used by said first motor.

33. The method of claim 29 further comprising engineering said design such that at least one of said tiltrotors is driven by a second motor.

34. The method of claim 33, further comprising including an additional plurality of motors in the design to individually power each rotor.

35. The method of claim 33, further comprising engineering the design such that a battery powers the first and second electric machines.

36. The method of claim 33 further including engineering said design such that at least one of said tiltrotors is additionally driven by a third motor.

37. The method of claim 29, further comprising engineering the design such that each rotor is powered.

38. The method of claim 29, wherein each rotor is driven by a separate motor, and further comprising including a battery in the design that powers all of the motors.

39. The method of claim 29, further comprising engineering the design such that tip speeds of all rotors when hovering are below 450 ft/sec.

40. The method of claim 29, further comprising engineering the design such that at least one rotor may continue hover operation over a continuous speed range between 60% and 100% of maximum RPM.

41. The method of claim 29, further comprising engineering the design such that at least one rotor can continue wing flight in a continuous speed range between 20% and 60% of maximum RPM.

42. The method of claim 29, further comprising engineering the design such that at least one rotor can apply a moment to control the pitch angle of the aircraft.

43. The method of claim 29, further comprising engineering the design such that the maximum lift to drag ratio for flight of the wing is at least 10.

44. The method of claim 29 further comprising engineering the rotor such that if the motor fails, the aircraft spins at a sustained descent rate of less than 1,000 ft/min.

45. The method of claim 29, further comprising engineering the wings and rotors such that a flight speed margin at the transition of the aircraft from VTOL flight to wing flight is not less than 20 KIAS.

46. The method of claim 29, further comprising engineering the wings and rotors such that the aircraft can maneuver at a maximum weight of 3g without loss of altitude or speed.

47. The method of claim 29, further comprising engineering the design such that the airfoil has an airfoil tip portion with a dihedral that actuates to move the airfoil tip to between 20-90 degrees.

48. The method of claim 47 further comprising engineering the design with a control system that adjusts the dihedral wing tip to reduce wing down loads at hover.

49. The method of claim 47 further comprising engineering the design with a control system that adjusts the dihedral wing tips to provide roll support for taxiing in crosswinds.

50. The method of claim 47 further comprising engineering the design with a control system that adjusts the dihedral wing tips to tether an aircraft.

51. The method of claim 29 further comprising engineering the design such that at least one rotor is an auxiliary rotor that is no more than 50% of the disk area of the first rotor and provides additional lift and pitch control when hovering.

52. An electric VTOL aerial vehicle capable of carrying at least 500 pounds of payload, comprising:

a machine body

A wing mechanically coupled to the fuselage, the wing sized and dimensioned to provide a wing load of no greater than 40psf, and a wing-on-board stall of no greater than 90 KIAS;

at least first and second main-shift rigid rotors configured to provide lift to the fuselage, wherein each of the first and second rotors is open and not embedded in the wing;

each rotor has a plurality of blades with independent blade control and has a flap stiffness that is no less than the fourth power of the product of rotor diameter in feet times 200 at 30% of the rotor radius measured from the center of rotation of the rotor, wherein the flap stiffness is in lbs-in²Counting;

each of the first and second rotors being driven by at least one motor, and the first and second main rotors being sized and dimensioned to collectively provide at least 70% lift at vertical takeoff; and the at least one motor is powered by at least a first power source.

53. The aircraft of claim 52 wherein each of the main rotors is configured to provide a paddle load below 10psf and to provide a hover power load above 8 lb/HP.

54. The aircraft of claim 53 further comprising a hub bearing, a gearbox, and a motor mounting fixture that are together configured as an integrated rotor drive system.

55. The aircraft of claim 54 wherein the at least a first main rotor is powered by three motors.

56. The aircraft of claim 55 further comprising a wing tip portion having an actuator that adjusts the wing tip to a dihedral angle between 20-90 degrees.

57. The aircraft of claim 56 further comprising a tail wing and/or a front wing, each of said tail wing and/or front wing having a lifting surface having an area between 10% and 100% of the area of said wing.

Technical Field

Background

Disclosure of Invention

The present subject matter provides apparatus, systems, and methods in which an electric vertical take-off and landing (eVTOL) vehicle is engineered to carry at least 500 pounds (about 227 kilograms) of objects using a reduced number (e.g., 2 to 4) of variable speed rigid (non-articulated) rotors.

Various objects, features, aspects and advantages of the present subject matter will become more apparent from the detailed description of preferred embodiments and the accompanying drawings in which like numerals represent like components.

Drawings

FIG. 1A is forthcoming

Conceptual diagram of the urban transportation market for proposed hybrid electric vertical take-off and landing (eVTOL) aircraft.

FIG. 1B is a prior art development and operation schedule for the aircraft of FIG. 1A.

FIGS. 2A and 2B are artist versus prior art 16 rotor Volocopter^TMThe reproduction of (1).

FIG. 3 is a drawing of artsHome-to-prior-art 8-rotor wing Ehang^TMThe reproduction of (1).

Figure 4 is artist versus prior art 8 rotor type airfbus^TMThe reproduction of (1).

Figure 5 is artist versus prior art 36 rotor Lilium^TMRev of eVTOL, where the rotor is tilted around the front and rear wings.

Figure 6 is an artist versus prior art 8 rotor Airbus^TMA3 Vahana^TMReproduction of

FIG. 7 is a schematic representation of a prior art Joby^TMThe 6-rotor eVTOL concept images of the Computational Fluid Dynamics (CFD) flow solution of the aircraft.

FIG. 8 is artist vs Aurora^TMeVTOL, a reproduction of the concept aircraft, uses eight lift rotors and a rear-facing propeller.

FIGS. 9A and 9B show artist pairing Terrafugia^TMThe concept of an eVTOL aerial vehicle.

FIG. 10 shows a prior art Airbus^TMA3 Vahana^TMThe motor of (2) is installed.

Fig. 11 is a schematic perspective view of a preferred VTOL aerial vehicle according to the inventive concepts herein.

FIG. 12 is a table of dimensions and parameters for the aircraft of FIG. 11.

Fig. 13A and 13B are schematic top and side dimensional views, respectively, of the aircraft of fig. 11.

FIG. 13C is a table depicting possible seating arrangements and calculated weights for the aircraft of FIG. 11.

Fig. 13D and 13E are schematic perspective views of the aircraft of fig. 11 with the doors and hatches open.

Fig. 13F is a schematic side view of an open rear gangway of the aircraft of fig. 11.

Fig. 13G is a schematic illustration of a side cross-section of the aircraft of fig. 11.

FIG. 14 is a schematic side view of a slotted flap that may be used with the aircraft of FIG. 11, the slotted flap being shown in four different positions.

Fig. 15A is a schematic perspective view of an outer wing fold feature that may be used with the aircraft of fig. 11.

Fig. 15B is a schematic front view of the outer wing of fig. 15A in a folded orientation in the aircraft of fig. 11.

FIG. 15C is a schematic top view of the aircraft of FIG. 11 with the outer wings in a folded orientation so that the aircraft may be installed in a projected circular tarmac of diameter 45'.

Figure 16A is a table depicting rotor geometry and characteristics calculated as a function of the dimensionless radial station of the aircraft of figure 11.

Figures 16B-16E are graphs depicting calculated beaming-to-flapwise bending stiffness, chordwise bending stiffness, torsional stiffness, and mass per unit length for proposed rotor blades of the aircraft of figure 11.

Fig. 17 is a graph depicting the lowest blade natural frequency calculated by the aircraft of fig. 11 at a collective control setting of zero degrees based on rotor speed.

Fig. 18 is a schematic illustration of five cross-sectional wing profiles of the rotor blades of the aircraft of fig. 11 at designated radial stations.

FIG. 19A is a first schematic perspective view of a drive system that may be used with the aircraft of FIG. 11 packaged in a streamlined nacelle at a location desired for wing flight.

FIG. 19B is a second schematic perspective view of a portion of a drive system that may be used with the aircraft of FIG. 11.

FIG. 20 is a graph depicting a calculated effect of motor speed on motor weight for the aircraft of FIG. 11.

FIG. 21A is a schematic vertical cross-sectional view of a preferred Independent Blade Control (IBC) configuration that may be used with the aircraft of FIG. 11.

FIG. 21b is a schematic vertical cross-sectional view of an alternative preferred Independent Blade Control (IBC) configuration that may be used with the aircraft of FIG. 11.

Fig. 22 is a schematic perspective view of the aircraft of fig. 11 with Independent Blade Control (IBC) actuators according to fig. 21A or 21B, and further with a four-bladed main rotor and a four-bladed auxiliary rotor.

FIG. 23A is a schematic cross-sectional view of a nacelle that may be used with the aircraft of FIG. 11, wherein batteries are disposed below the wing and inside the nacelle.

FIG. 23B is a schematic cross-sectional view of a nacelle and wing that may be used with the aircraft of FIG. 11, with a battery disposed within the wing.

Fig. 24A-24G are schematic perspective views of an alternative preferred VTOL aerial vehicle according to the inventive concepts herein. This version has no secondary rotors.

Fig. 25A and 25B are schematic vertical cross-sectional views of the alternative VTOL aerial vehicle of fig. 24A-24G depicting 3 rows of seats comparable to the 4-rotor configuration of fig. 13G.

Fig. 26 is a table of dimensions and parameters for the 2-rotor alternative VTOL aerial vehicle of fig. 24A-24G.

Detailed Description

The present subject matter provides apparatus, systems, and methods in which an electric vertical take-off and landing (eVTOL) vehicle is designed to carry at least 500 pounds (about 227 kilograms) of objects using a reduced number (2 to 4) of variable speed rigid (non-articulated) rotors (typically assembled into a main rotor and a secondary rotor). Whether main and auxiliary, the rotors are preferably tiltrotors, such that one or more of the rotors provides a substantial amount of lift (e.g., 70%, etc.) during rotor flight (e.g., vertical takeoff, etc.), and the rotors may be tilted to provide forward thrust (or air braking) during wing flight.

In some contemplated embodiments, each rotor may be driven by its own motor or motors, and in other contemplated embodiments, multiple rotors may be driven by a single motor. In a particularly preferred embodiment, a single rotor may be driven by three motors. It is also contemplated that different motors may be powered by different battery packs, or that multiple motors may be powered by a single battery pack.

The terms "battery" and "battery" are used interchangeably herein and refer to one or more electrochemical cells that generate electricity. The battery preferably utilizes lithium ion chemistry and has a specific energy density of about 100 kWh/lb. Other contemplated battery chemistries include lithium polymers and lithium metal.

Non-articulated rotors are preferred because in VTOL and wing cruise flight, the pitch of the aircraft can be controlled by varying the individual blade angles to apply a moment. Blade angle control is preferably achieved by independent blade control actuators, preferably mounted within their respective blades, axially on the pitch axis. An independent blade control system used on at least each of the first and second main rotors applies a differential collective pitch between blades on the rotors such that rotor thrust remains substantially constant while shaft torque increases above that required without the differential collective. Details can be found in pending provisional applications, 62/513930(Tigner) "Propeller or Rotor in axial Flight for Aerodynamic Braking purposes" and 62/513925(Tigner) "improving the agility of the Rotorcraft dynamic Response through the Use of independent blade control (Use of Industrial blade control to Engine Rotor turbine Power Response Quickness)", each of which is incorporated herein by reference in its entirety.

In a preferred embodiment, the aircraft has a main rotor and a secondary rotor. The main rotor includes a plurality of blades and a hub configured to provide a moment of at least 6% of the rotor's maximum lift time at the rotor radius, more preferably at least 9% of the rotor radius, and most preferably at least 12% of the rotor radius.

To achieve commercially viable flight durations, lift and other characteristics without more than four rotors and currently available battery technology, at least the main rotor needs to be relatively large. Thus, each main rotor is configured to provide a paddle load below 10psf and a hover power load above 8 lb/HP.

More preferably, each main rotor is configured to provide a paddle load below 6psf and to provide a hover power load above 10 lb/HP. Other contemplated aircraft embodiments have a power load below 8 lb/HP.

Furthermore, in order to achieve high rotor efficiency in rotor and wing flight, it is necessary to continuously operate the rotor over a wide range of rotor speeds (RPM), e.g., 20% to 100%, contemplated embodiments utilize the rotor designs disclosed in U.S. patents 6007298(Karem) "optimal speed rotorcraft" (OSR) and 6641365(Karem) "optimal speed tiltrotor" (OSTR).

Using the teachings in OSR and OSTR, the aircraft contemplated herein preferably achieves a flap stiffness per blade that is no less than 100, preferably 200, times the fourth power of the product of the rotor diameter in feet, where the flap stiffness in lbs-in2 is measured at 30% of the rotor radius measured from the center of rotation of the rotor.

Also, using the teachings in OSR and OSTR, each blade weight in pounds preferably does not exceed the product of rotor diameter in cubic feet times 0.004.

Embodiments having first and second main rotors are contemplated that include at least one optional first auxiliary rotor, each first auxiliary rotor having a paddle area that is no greater than 50% of each main rotor. In a more preferred embodiment, the paddle area of each first auxiliary rotor is no more than 40% of each main rotor. The auxiliary rotors are not all the same size.

The one or more auxiliary rotors are also preferably rigid (non-articulated) rotors configured to generate a pitching moment by varying the pitch of the individual blades. The at least first auxiliary rotor is advantageously configured to provide a maximum aircraft pitching moment that is no greater than the collective total aircraft pitching moment capability of the main rotor.

For each rotor, the rotating hub, corresponding hub bearings, gear box, and motor mounting fixture are all configured together as an integrated rotor drive system. In a preferred embodiment, each main rotor comprises three independently controlled motors connected to a single gearbox. Three independently controlled motors provide a safety benefit through redundancy, and in addition, this configuration is a lightweight solution for the high torque output required for variable speed rotors.

A preferred embodiment includes a wing carrying at least first and second rotors, each rotor disposed in a rotor assembly configured to tilt at least 90 ° relative to the wing. In particularly preferred embodiments, the respective motor or other motive device is configured to tilt with the rotor assembly. At least the main rotors are open, i.e. they are not constrained by the circumferential air conduction band when the rotors are tilted.

Like rotors, wings are relatively large relative to the weight of the aircraft and payload. For example, the wing is preferably sized and dimensioned such that the wing load is no higher than 40psf and the wing stall speed is no higher than 90 KIAS. It is particularly preferred that the wing is configured such that the wing load is no higher than 20psf and the wing stall speed is no higher than 50 KIAS. The preferred wing is also configured to provide a flight speed margin of no less than 20KIAS in the transition from full rotor airborne level flight to full wing airborne level flight, and to provide a wing cruise lift/drag ratio of no less than 10. Particularly preferred wings are further configured to provide a transitional airspeed margin of no less than 40 KIAS.

To further reduce the stall speed of the aircraft, the preferred wing is fitted with an actuated slotted flap. In a particularly preferred configuration, the flaps may be used to provide aircraft roll control. The wing is preferably configured with a wing tip portion with a control system and a motorized or other actuator to adjust the wing tip to a dihedral angle between 20-90 degrees to: (a) reducing the downward load of the wing during suspension; (b) providing roll support for taxiing during crosswind; (c) the aircraft is restrained.

The wings, rotors, and other components and features discussed herein are preferably designed so that the aircraft can fly at a maximum weight of 3g without losing altitude or speed, but the aircraft can still maintain a low sustained autorotation descent rate if the rotors fail. Preferred embodiments have a sustained spin-down rate of less than 1,000 ft/min.

In some embodiments, at least a first battery or other power source is provided in the wing. Also, in some embodiments, the landing gear extends from at least one of the fuselage and the wing.

In another preferred embodiment, at least a first battery or other power source is disposed in the main rotor nacelle.

Embodiments are envisaged having a tail wing and/or a front wing, each of which preferably has a lifting surface having an area of between 10% and 100% of the area of the wing.

Contemplated embodiments include manned aircraft and unmanned aircraft. Thus, a place of self may have a passenger cabin with at least one seat for a person to sit.

Electronic controls sufficient to fly the aircraft without an onboard human pilot are also contemplated.

Fig. 11 is a perspective view of a preferred VTOL aerial vehicle according to the inventive concepts herein. The aircraft has a wing 1101, a static nacelle 1102, a tilt nacelle 1103, a fuselage 1150, a tail surface 1130, and a first tilt rotor system 1110. One particularly preferred embodiment includes a first tilt assist rotor system 1140.

Rotor system 1110 includes rotor blades 1120. The rotor blades are of the rigid, hingeless type, including for example the blades described in us patent No.6641365 (Karem). The rotor systems collectively provide thrust as indicated by arrow 1113 and torque 1114. Moments and forces may be controlled by rotating the blades about feathered axes 1121 extending along the length of blades 1120. The pitch angle about the feather axis 1121 is represented by arrow 1122. The tip of the rotor blade follows a rotational trajectory represented by circle 1116. Rotor blade 1120 and tilt nacelle 1103 can be tilted about tilt axis 1112 along a path represented by arrow 1111. To illustrate the tiltrotor function, the right-hand nacelle is in the wing flight direction, while the left-hand nacelle is in the rotor flight direction. During typical operation, the nacelles will be in a similar orientation.

Wings 1101 transfer loads from the rotor system to fuselage 1150. Fuselage 1150 is designed to carry payload and passengers, and includes various systems including landing gear.

Fig. 12 is a table of dimensions and parameters of a preferred embodiment, where "+" denotes turbulence. The preferred embodiment described in this table is designed for a nominal payload of approximately 1100 pounds and a basic mission takeoff weight of 4767 pounds. The wing area is 250 square feet and the resulting wing load is 19.1 psf. Including the effects of rotor wash flow on the fuselage, the total paddle pan area is 849 square feet, giving a hover pan load of 6.62lb/ft 2.

Figures 13A and 13B show top and side views, respectively, of a dimensional view of a preferred embodiment consistent with the aircraft shown in figure 11. 1310 are main landing gear wheels attached to the fuselage 1150. 1311 are nose landing gear wheels attached to fuselage 1150. All previously numbered elements are as described above.

FIG. 13B is a side view showing 2 upwardly opening front doors 1312 of the preferred embodiment aircraft, like the 4 doors 1313 and 1314 and the luggage compartment 1315 of an automobile, with 3 possible configurations: a) an aerial taxi with 1 pilot and 4 passengers, b) home use to accommodate up to 8 passengers, and c) freight/emergency use with foldable back row seats and an optional gangway instead of a luggage compartment. The home use configuration has a 1,350 pound payload capacity intended to accommodate a family like a large SUV, the only difference being that the SUV has 4 doors as compared to a 4 door SUV, while the aircraft has 2 front doors, equal to 6 doors provided.

FIG. 13C is a table depicting the seating arrangement and calculated weight of the preferred aircraft.

A payload capacity of 1,350 lbs, loading flexibility required when loading 3 rows of seats, especially 400 lbs at the rear of the trunk or rear gangway, can result in c.g. broadening of the aircraft to 8.5 inches (13.2% of the wing mean aerodynamic chord). In rotor flight, the combination of powerful pitch control of the auxiliary rotor and the pitching moment of the rigid main rotor by the preferred embodiment can provide aircraft stability and control such wide c.g. transitions; in wing flight, the effective pitching control combination of the large tail wing lifting bin and the pitching moment of the rigid main rotor wing can provide the stability of the aircraft and control the wide C.G. transformation.

Figures 13D and 13E show schematic views of the aircraft with the preferred embodiment with the doors and hatches open.

Fig. 13F shows the aft gangway 1316 open. The nose landing gear 1311 is optionally of variable height, allowing the fuselage to be adjusted to ground angle, providing additional clearance at the gangway opening.

Fig. 13G shows a side view of a fuselage with a proposed seat arrangement consistent with fig. 13C. Electronics 1317 for later use are contemplated that enable the aircraft to fly without an onboard human pilot.

Aerodynamic design

The aircraft contemplated herein is designed for efficient vertical and cruise flight. Furthermore, such aircraft are designed to provide safe flight and perform well in transitional flight conditions between full wing flight and rotor flight (which are referred to as "transitions").

The rotor thrust required for vertical flight is approximately 10 times that required for efficient cruise flight. The preferred embodiment aircraft uses a variable speed rotor as described in U.S. patent 6641365(Karem) to efficiently transition from 100RPM in low speed wing flight to 460RPM at hover at 12,000 feet. The aerodynamic design of the rotor gives way to the optimal behaviour for hover and cruise flight, where the rotational speed range for hover and cruise flight in this rotor is 5: 1. it is desirable to have a combination of airfoil designs that have linear lift characteristics over various ranges of angle of attack, twist angle and chord distributions that balance vertical and cruise flight. Cross-section airfoil design and analysis tools (e.g., XFOIL) may be used to design and study airfoils that achieve desired characteristics. Rotor analysis software for cruise (e.g., XROTOR) and software for hover rotor performance (e.g., charm (cdi)) may be used to optimize rotor geometry to achieve desired performance characteristics. The resulting preferred rotor geometry is given in the table in figure 16A, while the airfoil section is given in figure 18.

For efficient cruise flight, lift resistance ratios at least as high as 10 are desired. By using Computational Fluid Dynamics (CFD) programs (e.g., STAR-CCM +), the shape can be analyzed and iteratively optimized based on practical considerations (e.g., propulsion and payload volume and structural requirements) to minimize fuselage and nacelle drag. The wing airfoil may be optimized using an airfoil tool such as the aforementioned XFOIL. Considerations related to wing optimization include compromises between cruise drag, vertical flight down-load, transitional maximum lift, and structural requirements.

As shown in the cross-sectional view of FIG. 14, a preferred method of increasing maximum lift during a transition without negatively impacting cruise resistance is slotted flaps. The map includes a plurality of flap yaw positions: maximum upward deflection (-8), no deflection (0), maximum CL (+22), and maximum downward deflection (+ 65). The flap 1401 rotates about a simple hinge 1404. When retracted (position 0), the flap 1401 causes minimal additional drag compared to the airfoil of a single element. When deployed to an optimal high lift angle (position +22), a slit is exposed that allows airflow from the underside of the first member 1402 through the flap member 1401. In a preferred embodiment, the flaps and slots are shaped to provide a linear lift response to a small deflection angle, thereby enabling the slotted flaps to also be precisely controlled as ailerons for aircraft roll control. The preferred slotted flap extends spanwise from the side of the fuselage to the tip with an interruption at the propulsion system nacelle. The flap is divided into a plurality of spanwise portions to reduce the stresses caused by wing deformation. The flap hinge is offset from the wing surface so that when the flap deflects downward, the slit opens. The flexible upper surface seal 1403 minimizes drag when the flap is in its retracted position.

The most critical flight safety phase in wing eVTOL flight is the transition from full rotor onboard to wing onboard with safe forward speed. This is especially important in high winds and low altitude urban environments. Unlike the prior art, the use of a large main rotor in combination with a large wing area provides a safe transition.

In the preferred embodiment (with 2 secondary rotors), the use of a large rotor results in paddle loading, resulting in low noise (350RPM, rotor tip mach number less than 0.35), efficient stable hover and rotor handling of 2g at 495 RPM. In combination with a large (250Ft ^2) wing with 22 degree slotted flaps, the aircraft was designed to provide a stall speed of 50KIAS at an aircraft weight of 4,767 pounds. At a rotor speed of 550RPM, the aircraft may be fully on-board the rotor (zero wing and tail lift), at a speed of 90KIAS, 40KIAS higher than the minimum wing flight speed, and may carry 2.5g of instantaneous rotor lift. At 90KIAS, the wing lift of the aircraft is 3.25 g. These large margins avoid most flight accidents caused by low lift and control accidents caused by flight and transition at low speeds in extreme weather.

Folding outer wing

Figure 15A depicts an outer flap fold feature. The outer wing 1501 folds about the hinge line 1503 relative to the inner wing 1502 and this folding action is controlled by a folding actuator (not shown). The flap actuator is designed to withstand flight and ground loads. At its wing tip there is a spring-suspended slide 1504, which spring-suspended cleats 1504 contact the ground when landing. The large wingspan and acceptable transition characteristics required for efficient flight make the aircraft sensitive to ground crosswinds and gusts. The wingtip cleats provide additional ground stability and safety for the aircraft. The wing tip also comprises fixing features (not shown) for fixing the aircraft when parked. In rotor flight, because the downward load is inversely proportional to the rotor-wing separation distance, the downward load due to rotor wash on the outer wing is reduced. Figure 15B shows a front view of the aircraft with the outer wings folded. In addition, folded aircraft can be accommodated in smaller parking areas. Fig. 15C shows a top view of the aircraft, which is adapted to park in a projected circular tarmac of 45 inches in diameter.

Blade design

Figure 16A shows rotor geometry and characteristics as a function of dimensionless radial station. Figure 18 shows a cross-sectional airfoil profile of a rotor blade at a given radial station.

Figure 16B depicts the flapwise, or normal to chord, bending stiffness of the rotor blade of the exemplary embodiment from radial station 0 at the root to radial station 1 at the tip. Fig. 16C depicts the lag or chordwise bending stiffness of the rotor blade of the exemplary embodiment from radial station 0 at the root to radial station 1 at the tip. Figure 16D depicts the torsional stiffness of the rotor blade of the exemplary embodiment from radial station 0 at the root to radial station 1 at the tip. Figure 16E depicts the mass per unit length of the rotor blade of the exemplary embodiment from radial station 0 at the root to radial station 1 at the tip. Fig. 16F depicts the chordwise cg position of the rotor blade of the exemplary embodiment relative to the blade pitch axis from radial station 0 at the root to radial station 1 at the tip.

FIGS. 16C-E depict the spanwise bending stiffness, chordwise bending stiffness, and torsional stiffness of the proposed embodiment blade. A high stiffness to mass ratio is required to operate the rotor over a wide range of rotor speeds without structural dynamics issues. The blade mass distribution of the proposed embodiment blade is shown in fig. 20F. To avoid aeroelastic instability, the chordwise center of mass of the rotor blade cannot be located further aft than the blade pitch or feather axis. As shown in fig. 16F, the rotor blades of the embodiments were found to exhibit no aeroelastic instability under operating conditions and to be balanced in terms of center of mass.

Fig. 17 depicts the lowest blade natural frequency as a function of rotor speed at a common setting of zero. The ray from the origin depicts the 1/rev, 2/rev … 10/rev harmonic frequency of the rotor. The operating speed range is marked on the horizontal axis. The stiffness and mass distributions of FIGS. 16A-F are the primary effects on the natural frequency of the rotor blade. The natural frequencies remain well separated from each other throughout the operating range. Due to the high stiffness and weight design, the natural frequency is much higher than a typical rotor blade. The first flap mode remains above the rotor's 3/rev excitation frequency throughout the operating range, while the first less damped hysteresis mode remains above the 4/rev excitation frequency. This separation above the main rotor excitation frequency of 3/rev allows operation over a wide range of rotor speeds without encountering excessive vibration loading or vibration due to resonance.

The described rotor blade design of desired characteristics may be iterated using rotor dynamics simulation and optimization software programs (e.g., CHARM and CAMRAD). Finite Element Analysis (FEA) software may be used for higher fidelity structural analysis, while CFD code may be used for higher fidelity aerodynamic analysis and improvement.

The preferred auxiliary rotor and blade design follows the same performance limitations as the smaller diameter main rotor.

Hub drive system

The drive system is enclosed in a streamlined nacelle, as shown in fig. 19A, at a location where the wing is required for flight. The axis of rotation of the rotor is shown as X-X and the direction of flight is shown as arrow A. By definition, tiltrotor aircraft require the thrust shaft of the rotor to be rotated from a horizontal flight condition to a vertical ascent condition. The angle is not less than 90 degrees and may be 105 degrees or more. The axis along which the front part of the nacelle (including all the drive elements) is tilted is shown as Y-Y.

The blade shank 1901 is mounted in a slide bearing receiving ring 1902, and the slide bearing receiving ring 1902 is bolted to a rotating hub 1903, the hub 1903 being supported on a large diameter bearing 1904. Three motors 1905 are symmetrically disposed about the hub center. One of the motors 1906 is shown in cross-section 1906. The output sun gear 1907 is driven by a sprag clutch 1908. The planet gears 1909 are mounted in a planet carrier 1910, which planet carrier 1910 is connected to an output pinion 1911. Three identical output pinions mesh with ring gear 1912. The hub loads are carried by the hub bearings through the intermediate structure 1913, and the intermediate structure 1913 is secured by bonding and riveting to the fuselage cell composite structure in the nacelle. The housing structure is attached to the nacelle at hinge point 1917, and a sloped actuation truss 1916 connects the two nacelle elements at actuator attachment bracket 1917. The electronic motor drive case 1918 is individually packaged for redundancy, with the link 1919 connected to the motor. The motor liquid cooling connections 1920 are shown, as well as an oil tank 1921. The alternative rotary tilt actuator 1922 shown is mounted on a transverse axis of rotation.

FIG. 19B depicts an alternative perspective view of a portion of the streamlined nacelle depicted in FIG. 19A. Fig. 19B provides a detailed view of the blade shank 1901 mounted in the slide bearing receiving ring 1902, and a detailed view of the bolted connection of the slide bearing receiving ring 1902 to the rotating hub 1903, the rotating hub 1903 being supported on the large diameter bearing 1904.

The entire rotor hub, including the blade slide bearings and the pitch actuation system coupled to the electric drive, forms an integral assembly. The system is shown as a three-bladed arrangement; other blade numbers are similarly installed. The four main loads addressed by the assembly from the rotating gantry to the nacelle structure are the blade flap load, the mast moment, the thrust or lift vector, and the drive torque. A large diameter moment-carrying bearing connects the rotating hub element to the nacelle structure. Large slow rotating rotors create drive conditions where the torque/speed characteristics of the rotor far exceed the capability of a direct drive motor. Fig. 20 shows a summary analysis of the effect of motor speed on motor weight, indicating that for constant power, a high speed motor with gear reduction can reduce weight with increasing gear ratio, and further reduce weight by multiplexing the motors. Since flight safety is critical, the multiplexed motors can provide complete electrical redundancy. By arranging a one-way clutch ("sprag" clutch) on each motor output shaft, a degree of mechanical redundancy may also be provided. The system shown has three motors but is suitable for a larger number of motors.

Typical applications for direct drive motors for eVTOL lift rotors have several advantages. It is simple and simplicity gives inherent reliability. Moreover, it does not add extra weight to the gearbox. However, the weight can also be reduced by increasing the RPM of the motor and gearing the output stage. By exchanging the torque at the RPM with a fixed power, the weight can be greatly reduced. Since the cooling capacity is limited as the size is reduced, the weight reduction is limited. Thus, to optimize weight, the choice between direct drive and gear motor depends on the required power output and the required RPM. At lower output RPMs, the gear drive has a weight advantage, while at higher RPMs the weight advantage of the direct drive is more pronounced. Figure 20 shows the weight of the motor and the drive force at 300HP rated output of the main rotor in a preferred embodiment. In this case, the weight breakpoint is near 2000 RPM. In the preferred embodiment, the main rotor RPM at high power hover is 400-.

Incorporating multiple motors in a design has several advantages. A) More motors can be more efficient in weight reduction. The heat dissipation capacity is proportional to the surface area of the motor, while the power (at a fixed RPM) is proportional to the volume of the motor, and thus to the weight of the motor. The larger the surface area to volume ratio, the better the cooling effect. In the case where the minimum weight is primarily determined by the cooling capacity, more motors may provide better weight efficiency because they have a higher surface area to volume ratio. B) Reliability can be improved by redundancy. Overall reliability may be improved by redundancy in configurations where acceptable output power may be maintained in the event of failure of one or more drive motors. However, the complexity of many motors can reduce reliability.

The higher the gear ratio, the more effective the weight reduction. Because the weight of the gearbox is driven to first order by the high torque output stage, the weight of the gearbox is independent of the gear ratio. As shown in fig. 20, the weight of the motor is inversely proportional to the RPM. However, as the size decreases, weight reduction is limited by a decrease in heat dissipation capacity. In addition, there are some practical limitations to motor RPM, including: holding the magnets at high centrifugal force (for this design type of motor), limits on the available bearing speed and on the electronic switching speed of the motor commutation.

In the present example, 20: the 1 gear ratio reduces weight while limiting the challenge of very high motor RPM. In other contemplated embodiments, the gearbox may have a 3: 1. 5: 1. 10: 1. 20: 1 or 30: 1, in the presence of a catalyst.

When the total gear ratio approaches or exceeds 20: 1, two stages of gear reduction are required to achieve the full weight reduction benefit of the high speed motor. Each motor is equipped with a planetary reduction set that drives a ring gear connected to the hub. All functions of the assembly are optimised to achieve a minimum weight, for example, using three drive pinions in mesh with a single large ring gear minimises the face width of the ring gear and thus saves material.

The motor, its driver electronics and the gearbox all need to be cooled. The preferred fluid for cooling the motor and electronics is water/glycol, and in addition a liquid-to-oil heat exchanger is used to cool the transmission oil. The transmission oil is contained in an oil pan at the lower end of the transmission housing.

The nacelle tilt system shown is a system containing 3 linear actuators, providing an aft of 60 degrees of the nacelle and a fwd of 55 degrees of the remaining actuators. An alternative system is the application of high torque rotary actuators operated by four bar linkages. The rotary actuator whose operational principle is described in detail is cited in the reference, (US7871033(Karem et al).

Independent blade control

In a preferred embodiment, Independent Blade Control (IBC) actuators 2101 are capable of accurately and independently controlling the trajectory of the rotor blades. By independently controlling blade angle, rotor torque and force can be controlled. Fig. 22 is similar to fig. 11, except that first tilt rotor system 2210 and first tilt rotor system 2240 in fig. 22 are for a 4-blade rotor application. The application of IBC actuation of a 4-blade rotor allows the wing to generate aircraft braking (rotor thrust) in flight without introducing large hub moments. All elements similarly numbered as in fig. 11 are described above. Details can be found in pending provisional applications, 62/513930(Tigner) "Propeller or Rotor in Axial Flight for aerodynamic braking purposes" and 62/513925(Tigner) "improving the agility of the Rotorcraft dynamic response through the Use of independent Blade Control (Use of Industrial Blade Control to aircraft Power response quickness)", each of which is incorporated herein by reference in its entirety.

Fig. 21A shows a preferred IBC configuration. The design approach is to position the electric actuator within the blade itself and to position the electric actuator such that the actuator and the blade slip axis are present simultaneously. Certain blade design conditions must be met to make the above approach feasible. The OSTR rotor blade should have the great advantage of having a high stiffness in terms of flap bending and lead-lag, which results in a blade root chord length and thickness that greatly exceeds conventional rotor blades. The resulting blade spar is hollow and of sufficient diameter to easily accommodate a cylindrical electric actuator. In combination with the reduction gearbox, it can be seen that the motor drive can rotationally connect the blades to the hub without mechanical linkages and that the electric drive can be precisely commanded and controlled as other flight control actuators. A generic term for such actuation is Independent Blade Control (IBC), which allows the use of completely new, optimized blade azimuth and pitch angle matrices. This has aerodynamic advantages.

In addition, military helicopters that engage in shipboard combat must be made more compact by folding. Folding prior art rotor blades and maintaining the integrity of the pitch link results in a complex arrangement of mechanical parts. The present invention eliminates this complexity. The only new requirement for actuator design inside the blade is that the cable wires are bent at a dog-ear. This requirement can be easily met at present.

Hollow blade spars 2101 are inserted into receiving holes in hub 2102, the hub 2102 supporting the blades by bearings 2103 on inner ring 2104, which inner ring 2104 is sealed by seals 2105. If folding is desired, the blade and hub portions rotate about hinge 2106. One motor stator 2107, or two motor stators 2107 and 2108, guide the rotation of the rotor 2109 by the tail bearing 2110 and the rotor bearing 2111. The encoder 2112 detects the rotor position and the blade angular position of the motor through the stationary core 2113 based on the static reference. The motor drives a gearbox 2114, which gearbox 2114 is secured to the blade root by fasteners 2115. The gearbox reaction torque is carried by the flexure coupling 2116, which is intended to isolate the gearbox from moment-induced deflections caused by blade flaps and lead-lag loads. The radial load due to the centrifugal load and the moment is carried by the tapered roller bearing 2117. Blade actuation torque is reacted through splines 2117, while centrifugal force is reacted through nuts 2119. The flexible electrical connection cable 2120 carries motor power and control force information from the slip ring 2121, which rotates about the hub rotation axis 2122 (shown by the XX axis) while the stationary portion of the slip ring is supported by the airframe structure 2123. The coolant flow and return line 2124 are supplied through a swivel 2125.

Fig. 21B shows an alternative component layout. If it is not necessary to fold the rotor blades and the blade taxi axis is tightly controlled with respect to the hub by means of rigid sliding bearings, the pitch actuators may be mounted on the hub instead of the blades. In this arrangement, an actuator assembly consisting of one or more motors, a reduction gearbox and the required sensors and connecting wires is connected to the hub. This connection is torsionally rigid, but is flexible in alignment and can withstand the flexure inherent in high load blades. The spline output driving disc is matched with a meshing spline in the paddle. Mounting a set of actuators together on an integral hub has practical system advantages in that the actuators can be electrically connected to a common driver, power supply and cooling path.

Cylindrical blade spar 2131 is supported in an outboard sliding bearing assembly consisting of bearing retainer ring 2132, outer race 2133, roller and cage 2134, inner race 2135 and seals 2136. The blade root is stabilized by an inner spacer 2137 and fixed by rivets 2138. The diaphragm is internally splined at 2139 for torque transfer from the flexible drive bellows 2140. This is the point of separation when the detached paddle is retracted onto the stationary actuator.

Split blade retention clips 2141 secure the inboard root fitting 2142 to the inboard sliding bearing outer race 2143. The tapered rollers and cage 2144 run on an inner ring 2145, and the inner ring 2145 is sealed by a seal 2146. Bearing preload is provided by a 2147 year disc spring 2147, the disc spring 2147 running over a thrust washer 2148.

The blade centrifugal forces are reacted by the actuator housing 2149, which is held in place by fastener assemblies 2150, which also secure the static core 2151 to the rotating hub assembly 2152. The static core carries both motor stator windings 2153 and position encoder 2154. Motor rotors 2155 are supported in journal and tail bearings 2156 and 2157 and drive a reduction gearbox 2158.

A tubular extender 2159 attached to the non-rotating fuselage structure carries a slip ring 2160 that provides current and control signals to the actuator via a fixed wire harness 2161.

Battery with a battery cell

A preferred battery mounting is shown in the nacelle cross-sectional view of fig. 23A. Batteries 2301 are disposed below wing 2302 and inside nacelle 2303. The direction of flight is indicated by the block arrow B. The rear volume of the nacelle contains enough volume to include the parts of the cooling system required by the electric propulsion system. Alternatively, as shown in fig. 23B, a battery 2311 may be housed in the wing structure 2312. The illustrated cell 2311 is smaller in cross-sectional size than the cell in fig. 23A, but has the same volume as the cell 2311 is surrounded by the long wings 2321. Thus, the nacelle 2313 is much smaller and the drag is also smaller than in the main preferred embodiment. Alternatively, a nacelle arrangement with an internal combustion engine and generator (hybrid) may be used to power electric motors and other aircraft systems and provide longer range.

Alternative arrangements

Fig. 24A-G show an alternative preferred embodiment without an auxiliary rotor. This preferred embodiment utilizes powerful pitch control of the main rotor 2401 in rotor flight and pitch control of the long control arm canard surface 2411 in wing flight, providing the same performance of the aircraft but with reduced aircraft control in transients and gusts and reduced containment levels of c.g. conversion and reduced payload versatility (cargo ramps and reduced luggage volume and weight) as compared to the preferred embodiment with auxiliary rotors and large tail wing area.

The alternative preferred embodiment features a fuselage 2421, wings 2431, rotor blades 2401 and canard 2411. The internal construction is similar to the main preferred embodiment. The fuselage has 3 rows of seats: a front row 2501, a middle row 2502, and a rear row 2503.

Although the cabin volume of the two-rotor alternative arrangement is comparable to that of an aircraft with four rotors, its wing drag is reduced by: a) wing area is reduced from 250Ft 2 to 140Ft 2, b) there is no tail, c) there is no auxiliary rotor nacelle, d) the fuselage is attached to the wing attachment behind the passenger cabin (lower frontal area), e) the choice of laminar fuselage flow is wide, and f) the cruise drag is reduced due to the reduction in cruise weight (due to lighter fuselage, smaller battery, predicted weight is also reduced by 817 pounds).

Fig. 25A and 25B show the inside profile of a 2-rotor alternative configuration, showing 3 rows of seats comparable to a 4-rotor configuration. Figure 26 is a table of dimensions and parameters for a two rotor alternative configuration.

Modifying an item

It will be apparent to those skilled in the art that many more modifications besides those already described are possible without departing from the inventive concepts herein. The inventive subject matter, therefore, is not to be restricted except in the spirit of the appended claims.